AIRCRAFT FUEL MANAGEMENT

Abstract

A method 2090 for determining at least one fuel characteristic of a fuel supplied to a gas turbine engine of an aircraft comprises carrying out 2092 a change in operation, the change in operation being effected by a controllable component of a propulsion system of which the gas turbine engine is a part, and being arranged to affect the operation of the gas turbine engine, detecting 2094 a response to the change in operation; and determining 2096 the at least one fuel characteristic based on the response to the change in operation. Figure 7

Description

Title of the invention: AIRCRAFT FUEL MANAGEMENT

[0001] This disclosure relates to aircraft propulsion systems and aircraft operating methods involving adaptations for fuels having different characteristics, and to methods for determining relevant fuel characteristics to enable the implementation of such methods.

[0002] The aviation industry expects the trend to be towards the use of fuels other than the traditional kerosene-based jet fuels generally used today. These fuels may have different fuel characteristics, for example, having one or both of these, a lower aromatic compound content and a lower sulfur content, compared to petroleum-based hydrocarbon fuels.

[0003] Thus, it is necessary to take into account the properties of fuel in light of the increased possibility of variation, and to adapt the control and management of aircraft propulsion systems and fuel supplies to these new fuels.

[0004] According to a first aspect, a method for controlling an aircraft propulsion system is provided, the propulsion system comprising: a gas turbine engine arranged to be fueled; and at least one steerable inlet guide vane (SIV). The method comprises:

[0005] obtaining at least one fuel characteristic of the fuel supplied to the gas turbine engine; and

[0006] the implementation of a programming change of at least one VIGV on the basis of at least one fuel characteristic obtained.

[0007] The at least one fuel characteristic may be or include at least one of the following: • the percentage of sustainable aviation fuel in the fuel; • the aromatic hydrocarbon content of the fuel; • the multi-aromatic hydrocarbon content of the fuel; • the percentage of nitrogen-containing species in the fuel; • the presence or percentage of a tracer species or trace element in the fuel (for example, a trace substance intrinsically present in the fuel which can vary from fuel to fuel and therefore be used to identify a fuel, and / or a substance deliberately added to act as a tracer); • the hydrogen to carbon ratio of the fuel; • the distribution of hydrocarbons from fuel; • the level of non-volatile particulate matter emissions during combustion (e.g. during combustion for a given combustion chamber design, under a given operating condition); • the naphthalene content of the fuel; • the sulfur content of the fuel; • the cycloparaffin content of the fuel; • the oxygen content of the fuel; • the thermal stability of the fuel; • the level of fuel coking; • an indication that the fuel is a fossil fuel; and • at least one of the following: density, viscosity, calorific value and heat capacity.

[0008] At least one fuel characteristic may be or include a fuel calorific value.

[0009] At least one fuel characteristic may be or include a fuel calorific value.

[0010] The step of carrying out a programming change of at least one VIGV may include the movement of at least one VIGV.

[0011] The step of implementing a programming change for at least one VIGV may include preventing or canceling a planned movement of at least one VIGV. For example, a closing step of a VIGV normally performed with a certain fuel, such as the widely used Jet A, at a certain point in the flight envelope may be canceled if the fuel being used has a higher calorific value than Jet A.

[0012] The propulsion system may include a plurality of fluidly separated fuel tanks containing different fuels so that the fuel supplied to the gas turbine engine can be changed in flight.

[0013] In such cases, the step of obtaining at least one fuel characteristic of the fuel supplied to the gas turbine engine may include determining an actual fuel or fuel mixture supplied to the gas turbine engine and obtaining the characteristic or characteristics for that fuel.

[0014] The step of obtaining at least one fuel characteristic can be repeated:

[0015] (i) at regular intervals;

[0016] (ii) each time the fuel or fuel mixture supplied to the gas turbine engine is changed; or

[0017] (iii) before each VIGV programming change.

[0018] The step of obtaining at least one fuel characteristic may include at least one of the following:

[0019] (i) the detection of at least one fuel characteristic, for example by physical and / or chemical detection processes, or the detection of parameters from which the fuel characteristic can be derived; and

[0020] (ii) retrieving at least one fuel characteristic, or data from which at least one fuel characteristic can be calculated, from a data storage.

[0021] At least one fuel characteristic may be or include a fuel calorific value - in such cases, the step of implementing a VIGV programming change may include opening at least one VIGV at takeoff by 1% of its range for each 1% increase in the fuel calorific value.

[0022] A linear or quasi-linear change in the VIGV angle can therefore be achieved with a change in calorific value.

[0023] At least one VIGV can have a full rotation range of 40°.

[0024] At least one fuel characteristic may be or include a ca capacity Fuel efficiency – in such cases, the step of implementing a VIGV programming change may involve opening at least one VIGV at takeoff by 0.5% of its range for a 30% increase in fuel calorific value. A linear or quasi-linear change in the VIGV angle can be achieved with the calorific value.

[0025] Opening at least one VIGV by 0.5% of its range for a 30% change in the fuel's calorific value can only be done up to a maximum additional opening of 5% of the VIGV's full range of motion. The at least one VIGV can have a full rotation range of 40°.

[0026] According to a second aspect, a propulsion system is provided for an aircraft, the propulsion system comprising:

[0027] a gas turbine engine arranged to be fueled and comprising:

[0028] a compressor; and

[0029] at least one steerable inlet guide vane - VIGV - through which / by which the airflow passes into the compressor;

[0030] and

[0031] a VIGV programming management device arranged for:

[0032] obtain at least one fuel characteristic of the fuel supplied to the gas turbine engine; and

[0033] perform a programming change of at least one VIGV on the basis of at least one fuel characteristic obtained.

[0034] The at least one fuel characteristic obtained may be or include a fuel calorific value.

[0035] The propulsion system may further include at least two fuel tanks containing different fuels so that the fuel supplied to the gas turbine engine can be changed in flight. In such cases, the VIGV programming management device can be configured to obtain at least one characteristic of the fuel currently supplied to the gas turbine engine:

[0036] (i) at regular intervals;

[0037] (ii) each time the fuel or fuel mixture supplied to the gas turbine engine is changed; and / or

[0038] (iii) before each VIGV programming change.

[0039] The propulsion system can be arranged to perform the process of the first aspect.

[0040] According to a third aspect, a method is provided for determining at least one fuel characteristic of a fuel supplied to a gas turbine engine of an aircraft, the gas turbine engine being part of an aircraft propulsion system. The method comprises:

[0041] the implementation of a change of operation to affect the operation of the gas turbine engine, the change of operation being carried out by a controllable component of the propulsion system;

[0042] the detection of a response to a change in operation; and

[0043] the determination of at least one fuel characteristic based on the response to change in operation.

[0044] The propulsion system can therefore be used to "carry out an experiment" to test the fuel, thus allowing one or more fuel characteristics to be determined based on the response of the gas turbine engine to the experiment.

[0045] Any suitable controllable component of the propulsion system can be used to bring about the change in operation. For example: • The propulsion system may include a thermal management system. The step of implementing a change in operation may include the use of, or consist of using, the thermal management system to change the temperature of the fuel entering a combustion chamber of the gas turbine engine, for example by adjusting the flow rates through one or more heat exchangers; • The propulsion system may include a fuel management system. The step of implementing an operating change may include or consist of changing the fuel flow rate and / or the fuel mixture; and / or • The propulsion system may include one or more guide vanes Adjustable Entry Gates (VIGV). The step of implementing a change of operation may include the movement of, or consist of moving, one or more VIGVs.

[0046] The response to the change in operation may include or consist of at least one of the following: i. a change in the output power of the gas turbine engine (for example, as indicated by an increase or decrease in shaft speed); ii. a change in fuel degradation or coking; iii. a change of at least one pressure in the engine; and / or iv. a change of at least one temperature in the engine.

[0047] The propulsion system may include at least one steerable inlet guide vane (VIGV). The step of implementing a change of operation may include changing or consist of changing the programming of the VIGV, for example by moving a VIGV, or by adjusting or canceling a planned movement of a VIGV.

[0048] The response to the change in operation of the VIGV programming may include or consist of at least one of the following: i. a change in the gas temperature at the inlet of a turbine in the gas turbine engine (for example, the Temperature at the Inlet of the High Pressure Turbine Rotor, T41); ii. a change in the temperature rise across a combustion chamber of the gas turbine engine (for example, reflected by the relationship between T30 and T41, where T30 is the temperature at the outlet of the high-pressure compressor); and iii. a change in the relationship between a total pressure at the compressor outlet - P30 - and a total pressure at the turbine rotor inlet - P41.

[0049] The propulsion system may include a plurality of fuel tanks. In such cases, the step of implementing a change of operation may include or consist of one and / or the other of the following: i. changing the tank from which the fuel is drawn; and ii. the change in the percentage of fuel taken from a particular tank (for example, switching to a different fuel mixture).

[0050] In such cases, the response to the change in operation may include or consist of one or more of the following: i. a change in the output power of the gas turbine engine; ii. a change in fuel degradation or coking; iii. a change in the formation of condensation trails; iv. a change in the relationship between a temperature at the compressor outlet and a temperature at the turbine rotor inlet; v. a change in the relationship between a total pressure at the compressor outlet and a total pressure at the turbine rotor inlet.

[0051] The propulsion system may include at least one air-oil heat exchanger. In such cases, the step of implementing a change in operation may include changing at least one of the air flow rates and the oil flow rates through the air-oil heat exchanger. The response to the change in operation may include a change in pressure in a fuel circuit of the gas turbine engine; for example, through a section of a pipe forming part of the fuel flow path, or through a pump, nozzle, or similar device.

[0052] At least one fuel characteristic may be or include at least one of the fuel characteristics listed above for the first aspect.

[0053] The determined fuel characteristic(s) produced by the process of this aspect can then be used to control the propulsion system and / or change a planned flight profile for a flight using the identified fuel, based on the determined fuel characteristic(s).

[0054] According to a fourth aspect, a propulsion system for an aircraft is provided, the propulsion system comprising:

[0055] a gas turbine engine;

[0056] a fuel tank arranged to contain fuel to power the gas turbine engine; and

[0057] a fuel composition monitoring device.

[0058] The fuel composition monitoring device is arranged to:

[0059] receive information concerning a change of operation, the change of operation being effected by a controllable component of the propulsion system and arranged to affect the operation of the gas turbine engine;

[0060] receive data corresponding to a response to the change in operation; and

[0061] determine one or more fuel characteristics of the fuel arranged to be supplied to the gas turbine engine based on the response to change in operation.

[0062] The propulsion system may further include one or more sensors arranged to detect a response to a change in operation. The sensor(s) may also be arranged to provide data concerning the response to the fuel composition monitoring device.

[0063] The sensor(s) may include one or more of a sensor temperature; and a pressure sensor. Multiple temperature and / or pressure sensors can be provided in different locations.

[0064] The propulsion system may further include one or more heat exchangers (for example, an air-oil heat exchanger, a fuel-oil heat exchanger, and / or a fuel-air heat exchanger, and possibly a plurality of one type of heat exchanger). The change in operation may include a change in at least one of the air flow rate, fuel flow rate, and oil flow rate through one or more heat exchangers. The propulsion system may further include one or more pressure sensors arranged to detect a pressure change in a fuel circuit of the gas turbine engine that may occur in response to such a change in operation; for example, a pressure change across a section of a pipe forming part of the fuel flow path, or across a pump, nozzle, or the like.It is understood that detecting a lack of pressure change despite a change in one or more heat exchange rates during fuel change can also be informative, and can help determine one or more fuel characteristics.

[0065] The gas turbine engine may include:

[0066] an engine core comprising a turbine, a compressor and a core shaft connecting the turbine to the compressor; and

[0067] 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.

[0068] The propulsion system may further include a flight profile adjustment device arranged to change the planned flight profile on the basis of the fuel or fuel characteristics.

[0069] The propulsion system may further include a propulsion system control device arranged to adjust the propulsion system control on the basis of the fuel or fuel characteristics.

[0070] The propulsion system can be arranged to implement the third aspect process.

[0071] According to a fifth aspect, a method is provided for determining at least one fuel characteristic of a fuel supplied to a gas turbine engine of an aircraft. The gas turbine engine is part of an aircraft propulsion system and comprises:

[0072] a combustion chamber arranged to burn fuel and having an outlet, and where a temperature at the outlet of the combustion chamber - T40 - is defined as an average flow temperature at the outlet of the combustion chamber in cruising conditions;

[0073] a turbine comprising a rotor having a leading edge and a trailing edge, and where a temperature at the inlet of the turbine rotor - T41 - is defined as an average flow temperature at the leading edge of the turbine rotor under cruising conditions; and

[0074] a compressor having an outlet, where a temperature at the outlet of the compressor - T30 - is defined as an average flow temperature at the outlet of the compressor under cruising conditions.

[0075] The process comprises:

[0076] changing the fuel supplied to the gas turbine engine; and

[0077] the determination of at least one fuel characteristic of the fuel based on a change of at least one of T30, T40 and T41.

[0078] The fuel characteristic(s) can be determined in terms of change for the fuel characteristic(s) relative to the previous fuel, and / or as absolute values.

[0079] The determination of at least one fuel characteristic can be based on a change in the relationship between T30 and one of T40 and T41. At least two of the temperatures can therefore be detected and used.

[0080] The relationship between the temperatures can be a difference between the temperatures. The difference between T30 and either of T40 and T41 can indicate a temperature rise across the combustion chamber.

[0081] The propulsion system may include at least one steerable inlet guide vane - VIGV.

[0082] No change in the position of VIGV can be made during the change of fuel, at least until after the determination of at least one fuel characteristic of the fuel (or at least until the capture of the data necessary for this determination to be made).

[0083] The change of fuel supplied to the gas turbine engine can be carried out during cruise.

[0084] The gas turbine engine may include multiple compressors. In such examples, the temperature at the compressor outlet can be defined as the temperature at the compressor outlet at the highest pressure.

[0085] The compressor may comprise at least one rotor, each rotor having a leading edge and a trailing edge. The temperature at the compressor outlet can be defined as the temperature at the axial position of the trailing edge of the rearmost rotor of the compressor.

[0086] The method may further include detecting a response to a change in fuel.

[0087] The at least one fuel characteristic may include at least one of the fuel characteristics listed above for the first aspect.

[0088] According to a sixth aspect, a method is provided for determining at least one characteristic of a fuel supplied to a gas turbine engine of an aircraft. The gas turbine engine is part of an aircraft propulsion system and comprises:

[0089] a combustion chamber arranged to burn fuel and having an outlet, and where a pressure at the outlet of the combustion chamber - P40 - is defined as the total pressure at the outlet of the combustion chamber under cruising conditions;

[0090] a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a pressure at the inlet of the turbine rotor - P41 - is defined as the total pressure at the leading edge of the turbine rotor under cruising conditions; and

[0091] a compressor having an outlet, where a pressure at the outlet of the compressor - P30 - is defined as the total pressure at the outlet of the compressor under cruising conditions.

[0092] The process comprises:

[0093] changing the fuel supplied to the gas turbine engine; and

[0094] the determination of at least one fuel characteristic of the fuel based on a change in at least one of P30, P40 and P41.

[0095] The determination can be carried out using at least two of the pressures, for example by evaluating a change in a relationship between P30 and one of P40 and P41.

[0096] The selected relationship between the pressures can be a pressure ratio.

[0097] Any characteristic as described with respect to the fifth aspect can be applied to this sixth aspect and, in some cases, the two can be used together - by examining both pressures and temperatures in order to determine or verify one or more fuel characteristics.

[0098] The gas turbine engine may include multiple compressors. In such examples, the pressure at the compressor outlet can be defined as the pressure at the outlet of the compressor at the highest pressure.

[0099] The compressor may comprise at least one rotor, each rotor having a leading edge and a trailing edge. The pressure at the compressor outlet can be defined as the pressure at the axial position of the trailing edge of the rearmost rotor of the compressor.

[0100] The determined fuel characteristic(s) produced by the fifth or sixth aspect process can then be used to control the propulsion system and / or change a planned flight profile, based on the determined fuel characteristic(s).

[0101] According to a seventh aspect, a propulsion system for an aircraft is provided, the propulsion system comprising:

[0102] a gas turbine engine comprising:

[0103] a combustion chamber arranged to burn fuel and having an outlet, and where a temperature at the outlet of the combustion chamber - T40 - is defined as an average flow temperature at the outlet of the combustion chamber under cruising conditions;

[0104] a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a temperature at the inlet of the turbine rotor - T41 - is defined as an average flux temperature at the leading edge of the turbine rotor under cruising conditions; and

[0105] a compressor having an outlet, where a temperature at the outlet of the compressor - T30 - is defined as an average flow temperature at the outlet of the compressor under cruising conditions;

[0106] a fuel tank arranged to contain fuel to power the gas turbine engine;

[0107] a fuel management device arranged to change the fuel supplied to the gas turbine engine; and

[0108] a fuel composition determination module arranged for:

[0109] receive data corresponding to a change in at least one of T30, T40 and T41; and

[0110] determine at least one fuel characteristic of the fuel based on the change in at least one temperature.

[0111] The fuel composition determination module can be arranged to receive data corresponding to at least two of the temperatures, and optionally to a change in the relationship between T30 and one of T40 and T41. The determination can be carried out on the basis of the change in the relationship between the temperatures.

[0112] The relationship between the temperatures can be a difference between the temperatures, the difference indicating a temperature rise across the combustion chamber.

[0113] The propulsion system may include at least two fuel tanks.

[0114] The propulsion system may further include at least one sensor arranged for provide data corresponding to one or more of T30, T40 and T41.

[0115] The propulsion system can be arranged to perform the process of the fifth and / or sixth aspect.

[0116] According to an eighth aspect, a propulsion system for an aircraft is provided, the propulsion system comprising:

[0117] a gas turbine engine comprising:

[0118] a combustion chamber arranged to burn fuel and having an outlet, and where a pressure at the outlet of the combustion chamber - P40 - is defined as the total flow pressure at the outlet of the combustion chamber under cruising conditions;

[0119] a turbine comprising a rotor having a leading edge and a trailing edge, and wherein a pressure at the inlet of the turbine rotor - P41 - is defined as the total flow pressure at the leading edge of the turbine rotor under cruising conditions; and

[0120] a compressor having an outlet, where a pressure at the outlet of the compressor - P30 - is defined as the total flow pressure at the outlet of the compressor under cruising conditions;

[0121] a fuel tank arranged to contain fuel to power the gas turbine engine;

[0122] a fuel management device arranged to change the fuel supplied to the gas turbine engine; and

[0123] a fuel composition determination module arranged for:

[0124] receive data corresponding to a change in a relationship between P30 and one of P40 and P41; and

[0125] determine at least one fuel characteristic of the fuel based on the change in the relationship between pressures.

[0126] The fuel composition determination module can be arranged to receive data corresponding to at least two of the pressures, and optionally to a change in the relationship between P30 and one of P40 and P41. The determination can be carried out on the basis of the change in the relationship between the pressures.

[0127] The propulsion system may include at least two fuel tanks.

[0128] The propulsion system may further include at least one sensor arranged for provide data corresponding to one or more of P30, P40 and P41.

[0129] The propulsion system of the seventh or eighth aspect may include a flight profile adjustment device arranged to change a planned flight profile for a flight of the aircraft based on the fuel or fuel characteristics.

[0130] The propulsion system of the seventh or eighth aspect may include a propulsion system control device arranged to adjust the propulsion system control on the basis of the fuel or fuel characteristics.

[0131] The propulsion system of the seventh or eighth aspect can be used to implement the process of the fifth and / or sixth aspect.

[0132] In the case where the gas turbine engine is an unshrouded rotor engine or a tur- In a gas turbine engine, the propeller can comprise two stages of counter-rotating propellers attached to and driven by a free-spinning turbine via a shaft. The propellers can 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 can comprise one propeller stage and a guide vane stage configured downstream of the propeller stage. The guide vane stage can have a variable pitch. Consequently, high-pressure, intermediate-pressure, and free-spinning turbines can drive high-pressure and medium-pressure compressors and propellers, respectively, via suitable interconnecting shafts. Thus, the propellers can provide the majority of the propulsive thrust.

[0133] When the gas turbine engine is an unshod rotor engine or a turboprop, one or more of the propeller stages may be driven by a gearbox of the type described.

[0134] The arrangements of this disclosure may be particularly, but not exclusively, advantageous for blowers that are driven by a gearbox. Accordingly, the gas turbine engine may include a gearbox that receives an input from the core shaft and delivers an output drive to the blower such that it drives the blower at a rotational speed lower than that of the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a gear and / or a straight shaft(s). The core shaft may rigidly connect the turbine and the compressor, so that the turbine and the compressor rotate at the same speed (with the blower rotating at a lower speed).

[0135]

[0136] 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 core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

[0137] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example, receiving directly, for example via a generally annular conduit) a flow from the first compressor.

[0138] The gearbox can be arranged to be driven by the core shaft that is configured to rotate (for example, during operation) at the lowest rotational speed (for example, the first core shaft in the example above). For example, the gearbox can be arranged to be driven only by the core shaft that is configured to rotate (for example, during operation) at the lowest rotational speed (for example, being only the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox can be arranged to be driven by any one or more shafts, for example, the first and / or the second shaft(s) in the example above.

[0139] The gearbox can be a reduction gear (in the sense that the output to the blower has a rotational speed lower than the input to the core shaft). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "star" gearbox, as described in more detail elsewhere here. The gearbox can have any desired gear ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or from 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 values ​​of the values ​​in the preceding sentence.Purely as an example, the gearbox may be a "star" gearbox with a ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

[0140] In any gas turbine engine as described and / or claimed herein, a fuel of a given composition or mixture is supplied to a combustion chamber, which may be located axially downstream of the fan and compressor(s). For example, the combustion chamber may be directly downstream (e.g., at the outlet) of the second compressor, when a second compressor is provided. As a further example, the outlet flow to the combustion chamber may be supplied at the inlet of the second turbine, when a second turbine is provided. The combustion chamber may be located upstream of the turbine(s).

[0141] The compressor or each compressor (for example, the first compressor and the second compressor as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades. (in the sense that their angle of incidence can be variable). The row of rotor blades and the row of stator blades can be axially offset from each other.

[0142] The turbine or turbines (for example, the first and second turbines as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades may be axially offset from each other.

[0143] Each fan blade can be defined as having a radial span extending from a foot (or hub) at a radially inward gas-washed location, or a 0% span position, to a tip at a 100% span 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) one of the following values: 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 limited by any two values ​​from the previous sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range from 0.28 to 0.32. These ratios can commonly be referred to as the hub / tip ratio.The radius at the hub and the radius at the tip can both be measured at the leading edge (or axially furthest point) of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e., the portion radially outside any platform.

[0144] The radius of the blower can be measured between the motor axis 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) one of the following values: 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 limited by any two values ​​from the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, within the range from 240 cm to 280 cm or from 330 cm to 380 cm.

[0145]

[0146] The rotational speed of the blower may vary during use. In general, the The rotational speed is lower for fans with a larger diameter. Purely by way of non-limiting example, the fan speed under cruising conditions may be less than 2500 rpm, for example, less than 2300 rpm. As a further non-limiting example, the fan speed under cruising conditions for an engine with a fan diameter in the range of 220 cm to 300 cm (for example, 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of 1700 rpm to 2500 rpm, for example, in the range of 1800 rpm to 2300 rpm, or for example, in the range of 1900 rpm to 2100 rpm.Purely as a further non-limiting example, the fan speed under cruising conditions for an engine with a fan diameter in the range of 330 cm to 380 cm can be in the range of 1200 rpm to 2000 rpm, for example in the range of 1300 rpm to 1800 rpm, for example in the range of 1400 rpm to 1800 rpm.

[0147] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the fan blade tip to move with a velocity Utip. The work done by the fan blades 13 on the flow results in an enthalpy increase dH of the flow. A fan tip load can be defined as dH / Utip², where dH is the enthalpy increase (e.g., the mean enthalpy increase 1-D) across the fan and Utip is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity).The fan peak load under cruise conditions can be greater than (or on the order of) one of the following values: 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 fan peak load can be within an inclusive range bounded by any two values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range from 0.28 to 0.31, or from 0.29 to 0.3.

[0148] Gas turbine engines according to this disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the stream through the bypass duct to the mass flow rate of the stream through the core under cruising conditions. In some arrangements, the bypass ratio may be greater than (or on the order of) any of the following values: 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 bypass ratio may be within an inclusive range limited by any two values ​​of the values ​​in the preceding sentence (i.e., the values (may form upper or lower limits), for example, in the range from 12 to 16, from 13 to 15, or from 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially external to the base motor. The radially external surface of the bypass duct may be defined by a nacelle and / or a blower housing.

[0149] 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 upstream of the fan to the stagnation pressure at the outlet of the highest-pressure compressor (before entering the combustion chamber). 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) one of the following values: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range limited by any two values ​​from the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range from 50 to 70.

[0150]

[0151] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, the specific thrust may depend, for a given thrust condition, on the specific composition of the fuel 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) one of the following values: 110 Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s, or 80 Nkg⁻¹s. The specific thrust can be within an inclusive range limited by any two values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range from 80 Nkg⁻¹s to 100 Nkg⁻¹s, or from 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines can be particularly efficient compared to conventional gas turbine engines.

[0152] 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 on the order of) one of the following values: 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 limited by any two values ​​from the preceding sentence (i.e., the values ​​may form upper or lower bounds). Purely as an 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, for example from 350 kN to 400 kN. The thrust mentioned above may be the maximum net thrust under normal atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine being stationary.

[0153] During operation, the flow temperature at the inlet to the high-pressure turbine can be particularly high. This temperature, which can be called TET, can be measured at the outlet to the combustion chamber, for example, immediately upstream of the first turbine blade, which can itself be called the turbine distributor blade. In some examples, the TET temperature may depend, for a given thrust condition, on the specific composition of the fuel supplied to the combustion chamber. During cruise, the TET temperature can be at least (or on the order of) one of the following values: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K, or 1650 K. The TET temperature during cruise can be within an inclusive range limited by any two values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds).The maximum TET temperature during engine operation can be, for example, at least (or on the order of) one of the following values: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET temperature can be within an inclusive range limited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range from 1800K to 1950K. The maximum TET temperature can occur, for example, under a high-thrust condition, such as a maximum takeoff time (MTO).

[0154] A fan blade and / or an airfoil portion of a fan blade described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be made at least partially from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as carbon fiber. By way of 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 or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may comprise at least two regions made using different materials.For example, a fan blade may have a protective leading edge, which can be made using a material more resistant to impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge could, for example, be made using titanium or a titanium-based alloy. Thus, purely as an example, the blade of... The blower may have a carbon fiber or aluminum (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0155] 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 attached to the central portion in any desired manner. For example, each blower blade may include a fastening element that can engage in a corresponding slot in the hub (or a disc). Purely by way of example, such a fastening element may be in the form of a dovetail that can be inserted into and / or engage in a corresponding slot in the hub / disk in order to secure the blower blade to the hub / disk. As a further example, the blower blades may be formed as a single unit with a central portion. Such an arrangement may be called a bladed disc or bladed ring.Any suitable process can be used to manufacture such a bladed disc or bladed ring. For example, at least part of the blower blades can be machined from a block and / or at least part of the blower blades can be attached to the hub / disc by welding, such as linear friction welding.

[0156] The gas turbine engines described and / or claimed herein may or may not be equipped with a variable area nozzle (VAN). Such a variable area nozzle can allow the outlet area of ​​the bypass duct to be varied during operation. The general principles of this disclosure may apply to engines with or without a VAN.

[0157] The blower of a gas turbine as described and / or claimed herein may have any desired number of blower blades, for example 14, 16, 18, 20, 22, 24 or 26 blower blades.

[0158] As used herein, the terms idle, taxiing, takeoff, climb, cruise, descent, approach, and landing have their conventional 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 would immediately recognize each term as referring to a phase of engine operation within the context of a given mission of an aircraft to which the gas turbine engine is designed to be attached.

[0159] 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 it is necessary for the engine to be running. At idle, the engine can produce between 3% and 9% of available engine thrust. In other examples, the engine can produce between 5% and 8% of available thrust. In still other 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 other examples, the engine can produce between 6% and 12% of available thrust. In still other 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. In an initial stage of the takeoff phase, the aircraft can be propelled while in contact with the ground. In a later stage of the takeoff phase, the aircraft can be propelled while not in contact with the ground. During takeoff, the engine can produce between 90% and 100% of available thrust. In other examples, the engine can produce between 95% and 100% of available thrust. In still other examples, the engine can produce 100% of available thrust.

[0160] The climb can refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. During the climb, the engine can produce between 75% and 100% of available thrust. In other examples, the engine can produce between 80% and 95% of available thrust. In still other examples, the engine can produce between 85% and 90% of available thrust. In this respect, the climb can refer to a phase of operation in an aircraft's flight cycle between takeoff and arrival under cruise conditions. Alternatively, the climb can refer to a nominal point in an aircraft's flight cycle between takeoff and landing, where a relative increase in altitude is required, which may necessitate an additional thrust demand from the engine.

[0161] As used herein, cruise conditions have the conventional 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 as denoting the engine's operating point at mid-cruise of a given mission (which in the industry may be called an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this respect, mid-cruise is the point in an aircraft's flight cycle where 50% of the total fuel burned between the end of the climb and the beginning of the descent has been burned (which may be approximated by the midpoint—in terms of time and / or distance—between the end of the climb and the beginning of the descent).Cruise conditions thus define an operating point for the gas turbine engine that provides thrust sufficient to ensure steady-state operation (i.e., maintaining a constant altitude and Mach number) at mid-cruise speed for the aircraft to which it is designed to be attached, taking into account the number of engines planned for that aircraft. For example, when an engine is designed... to be attached to an aircraft equipped with two engines of the same type, under cruise conditions the engine provides half of the total thrust that would be required for the steady-state operation of that aircraft at mid-cruise.

[0162] In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (necessary to ensure - in combination with any other engine on the aircraft - the steady-state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) under mid-cruise atmospheric conditions (defined by the International Standard Atmosphere in accordance with 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 thus the engine's operating point under cruise conditions is clearly defined.

[0163] Purely by way of example, the forward speed under cruise conditions can be any point in the range from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example in the order of Mach 0.8, in the order of Mach 0.85, or in the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

[0164] Purely by way of example, the cruising conditions may correspond to normal 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 (about 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 (about 35,000 ft) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range of 10,900 m to 11,100 m, for example on the order of 11000 m. Cruising conditions can correspond to normal atmospheric conditions at any given altitude within these ranges.

[0165] Purely by way of example, cruise conditions may correspond to an engine operating point that provides a known required thrust level (for example, a value in the range of 30 kN to 35 kN) at a Mach number of 0.8 and normal atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Purely by way of further example, cruise conditions may correspond to a point engine operating 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 advance of 0.85 and normal atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 feet (10668 m).

[0166] During operation, a gas turbine engine described and / or claimed herein can operate under the cruise conditions defined elsewhere herein. Such cruise conditions can be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one gas turbine engine (e.g., 2 or 4) can be mounted to provide propulsive thrust.

[0167] Furthermore, a person skilled in the art would immediately recognize descent and / or approach as referring to an operating phase in an aircraft's flight cycle between cruise and landing. During descent and / or approach, the engine may produce between 20% and 50% of available thrust. In other examples, the engine may produce between 25% and 40% of available thrust. In still other examples, the engine may produce between 30% and 35% of available thrust. Alternatively, descent may refer to a nominal point in an aircraft's flight cycle between takeoff and landing, where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

[0168] According to one aspect, an aircraft is provided comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

[0169] According to one aspect, a method of operating a gas turbine engine as described and / or claimed herein is provided. Operation can be carried out under cruising conditions as defined elsewhere herein (for example, in terms of thrust, atmospheric conditions and Mach number).

[0170] According to one aspect, a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein is envisaged. Operation according to this aspect may include (or may be) mid-cruise operation of the aircraft, as defined elsewhere herein.

[0171] A person skilled in the art will understand that, except as otherwise provided, a feature or parameter described in relation to any of the above aspects may be applied to any other aspect. Furthermore, except as otherwise provided, any feature or parameter described herein may be applied to any aspect and / or combined with any other feature or parameter described herein.

[0172]

[0173] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0174]

[0175] [Fig. 1] is a cross-sectional side view of a gas turbine engine;

[0176] [Fig.2] is a close-sectional side view of an upstream part of an engine gas turbine;

[0177] [Fig.3] is a partially cutaway view of a gearbox for a gas turbine engine;

[0178] [Fig.4] is a schematic view of VIGV through a compressor inlet of a gas turbine engine;

[0179] [Fig.5] is a schematic representation of a method for controlling an aircraft propulsion system;

[0180] [Fig.6] is a schematic view of an aircraft comprising a fuel composition determination module;

[0181] [Fig.7] is a schematic representation of a method for determining fuel characteristics;

[0182] [Fig.8] is a schematic view of an aircraft fuel composition monitoring system, in the context of a fuel supply line and an onboard tank, for use as a fuel composition determination module;

[0183] [Fig.9] is a schematic representation of a fuel characteristic determination method different from that shown in [Fig.7]; and

[0184] [Fig. 10] is a schematic representation of a propulsion system comprising an active fuel management system.

[0185] The [Fig. Figure 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 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 fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0186] During operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where a com Additional pressure is applied. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. 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 suitable interconnecting shaft 27. The blower 23 typically provides the majority of the propulsive thrust. The planetary gear box 30 is a reduction unit.

[0187] An exemplary arrangement for a gas turbine engine with a geared blower 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 planet gear, 28 of the epicyclic gear arrangement 30. A plurality of planet gears 32, coupled together by a planet carrier 34, are located radially outward from the planet gear 28 and mesh with it. The planet carrier 34 forces the planet gears 32 to precess around the planet gear 28 in synchronous rotation while allowing each planet gear 32 to rotate about its own axis. The satellite carrier 34 is coupled via connecting rods 36 to the blower 23 in order to drive its rotation around the motor shaft 9.A ring or crown 38 which is coupled, via connecting rods 40, to a stationary support structure 24 is located radially outwards from the planetary gears 32 and meshes with them.

[0188]

[0189] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" as used herein may be interpreted as referring respectively to the lowest-pressure turbine stages and the lowest-pressure compressor stages (i.e., not including the blower 23) and / or the turbine and compressor stages that are connected to each other by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the blower 23). In some publications, the terms "low-pressure turbine" and "low-pressure compressor" referred to herein may also be called "intermediate-pressure turbine" and "intermediate-pressure compressor." When such alternative nomenclature is used, the blower 23 may be called the first compression stage or the lowest-pressure compression stage.

[0190] The epicyclic gear box 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 parts of the The teeth are illustrated in [Fig. 3]. There are four planetary gears 32 shown, although it is obvious to those skilled in the art that more or fewer planetary gears 32 may be foreseen within the scope of the claimed invention. Practical applications of a planetary epicyclic gear box 30 generally include at least three planetary gears 32.

[0191]

[0192] The epicyclic gear box 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in the sense that the planet carrier 34 is coupled to an output shaft via connecting rods 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gear box 30 may be used. As a further example, the epicyclic gear box 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear (or ring) 38 permitted to rotate. In such an arrangement, the blower 23 is driven by the ring gear 38. As a further alternative example, the gear box 30 may be a differential gear box in which both the ring gear 38 and the planet carrier 34 are permitted to rotate.

[0193]

[0194] It will be understood that the arrangement shown in Figures 2 and 3 is by way of example only, and that various variants fall within the scope of this disclosure. 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. By way of further example, the connections (such as the connecting rods 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 bearing arrangement between the 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 housing) may be used, and disclosure 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 connecting rods and the bearing locations would typically be different from those shown as an example in [Fig. 2].

[0195]

[0196] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (e.g. star or planetary), support structures, input and output shaft arrangement and bearing locations.

[0197]

[0198] Optionally, the gearbox may drive additional and / or alternative components (for example, the intermediate pressure compressor and / or a pre-compressor).

[0199]

[0200] Other gas turbine engines to which this disclosure may apply 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. By way of further example, the gas turbine engine shown in [Fig. 1] has a split-flow nozzle 18, 20, i.e., the flow through the bypass duct 22 has its own nozzle 18 which is separate and radially outside the basic engine nozzle 20. However, this is not limiting, and any aspect of this disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be called a mixed-flow nozzle.One or both nozzles (whether a mixed-flow or split-flow nozzle) may have a fixed or variable surface area.

[0201] Although the example described relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as, for example, an unshod rotor engine (where the fan stage is not enclosed in a nacelle) or a turboprop. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0202] The geometry of the gas turbine engine 10, and of its components, is defined by a conventional axis system, comprising an axial direction (which is aligned with the 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.

[0203] The fuel F supplied to the combustion equipment 16 may include a fossil hydrocarbon-based 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 aromatic compounds. 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 given, fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganic matter, and metals.

[0204] The functional performance of a given composition, or of a fuel mixture to be used in a given mission, can be defined, at less so, in part, by the fuel's ability to sustain the Brayton cycle of the gas turbine engine. 10. Parameters defining functional performance can include, for example, specific energy; energy density; thermal stability; and emissions, including particulate matter. A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, can at least partially reduce the takeoff weight, thus potentially ensuring 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 takeoff fuel volume, which can be particularly important for volume-limited missions or military operations involving refueling.Relatively higher thermal stability (i.e., the inhibition of fuel degradation or coking under thermal stress) can allow the fuel to withstand high temperatures in the engine and fuel injectors, potentially resulting in relative improvements in combustion efficiency. Reduced emissions, including particulate matter, can help minimize contrail formation, thus reducing the environmental impact of a given mission. Other fuel properties can also be critical to functional performance.For example, a relatively lower freezing point (°C) can allow long-range missions to optimize flight profiles; minimum concentrations of aromatic compounds (%) 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 compound content; and, a maximum surface tension (mN / m) can ensure sufficient spray break and atomization of the fuel.

[0205] The ratio of hydrogen to carbon atoms in a molecule can influence the specific energy of a given fuel composition or mixture. Fuels with higher hydrogen-to-carbon ratios may have higher specific energies in the absence of bonding constraints. For example, fossil hydrocarbon fuels may comprise molecules with approximately 7 to 18 carbon atoms, with a significant portion of a given composition consisting of molecules with 9 to 15 carbon atoms, and an average of 12 carbon atoms.

[0206] ASTM International (ASTM) Standard D7566, Standard Specification for Jet Fuel Containing Synthesized Hydrocarbons (ASTM 2019c), approves a number of sustainable aviation fuel blends comprising between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fuels with based on fossil hydrocarbons, such as kerosene), with other compositions pending approval. However, the aviation industry anticipates that sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) will eventually be approved for use.

[0207] Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cyclo-alkanes, and aromatic compounds, and may be produced, for example, from one or more of a synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise a lower content of aromatic compounds and / or sulfur compared to fossil hydrocarbon fuels. Alternatively, sustainable aviation fuels may comprise a higher content of iso-alkanes and / or cyclo-alkanes compared to fossil hydrocarbon fuels. Thus, in some examples, sustainable aviation fuels may have a density between 90% and 98% of that of kerosene and / or a calorific value between 101% and 105% of that of kerosene.

[0208] Due, at least in part, to the molecular structure of sustainable aviation fuels, sustainable aviation fuels can provide advantages including, for example, one or more of the following: higher energy density; higher specific energy; higher specific heat capacity; higher thermal stability; higher lubricating power; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; and lower CO2 emissions, compared to fossil hydrocarbon-based fuels (e.g., when burned in combustion equipment 16). Consequently, compared to fossil hydrocarbon-based fuels, such as kerosene, sustainable aviation fuels can result in a relative decrease in specific fuel consumption and / or a relative decrease in maintenance costs.

[0209] As used herein, T30, T40, T41, P30, P40 and P41, and all other numbered pressures and temperatures, are defined using the station numbering listed in SAE AS755, in particular: • P30 = Total Pressure at the Outlet of the High Pressure Compressor (HPC); • T30 = Temperature at the HPC outlet; • P40 = Total Pressure at the Outlet of the Combustion Chamber; • T40 = Temperature at the Combustion Chamber Outlet; • P41 = Total Pressure at the Inlet of the High Pressure Turbine (HPT) Rotor • T41 = Temperature at the Inlet of the HPT Rotor.

[0210] As shown in [Fig. 6], an aircraft 1 may include multiple 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 1 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 (all the fuel being instead stored in the wings of the aircraft) – it will be understood that many different tank configurations are envisaged and that the illustrated examples are provided for ease of description and are not intended to be limiting.

[0211] Figure 6 shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply circuit 3 on board the aircraft. The fuel supply circuit 3 in the illustrated example comprises a single fuel source. For the purposes of this application, the term "fuel source" is understood to mean 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 separate fuel source; that is, the first fuel source may contain a first fuel having a different characteristic or characteristics from a second fuel contained in a 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).

[0212] In the present example, the first fuel source comprises a central fuel tank 50, located primarily in the aircraft fuselage, 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 the tanks 50, 53 are fluidly interconnected in the illustrated example, thus forming a single fuel source. Each of the central fuel tank and the wing fuel tanks may comprise a plurality of fluidly interconnected fuel tanks.

[0213] In another example, the wing fuel tanks 53a, 53b may not be in fluidic communication with the central tank 50, thus forming a second, separate fuel source. For balancing purposes, one or more fuel tanks in the left wing may be in fluidic communication with one or more fuel tanks in the right wing. This may be achieved via a central fuel tank 50 (if this tank is not part of the other fuel source), or by bypassing the fuel tank(s). central / centrals, or both (for maximum flexibility and security).

[0214] In another example, the first fuel source comprises wing fuel tanks 53 and a center fuel tank 50, while a second fuel source comprises another separate center fuel tank (not shown). A fluidic interconnection between wing fuel tanks 53 and the center fuel tank 50 of the first fuel source may be provided for balancing the aircraft 1.

[0215] In certain examples, the distribution of fuel tanks 50, 53 available on the aircraft may be constrained so that the first and second fuel sources are each substantially symmetrical with respect to the aircraft's axis. In cases where an asymmetrical distribution of fuel tanks is permitted, a suitable means of fuel transfer may be provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source so that the position of the aircraft's center of mass can be maintained within acceptable lateral limits throughout the flight.

[0216] An aircraft 1 can be refueled by connecting a fuel storage container 60, such as that supplied by an airport tanker truck, or a permanent pipeline, to a fuel line connection port 62 of the aircraft, via a fuel line 61. A desired quantity of fuel can be transferred from the fuel storage container 60 to the 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 fuel line connection ports 62 may be provided instead of one, and / or valves may be used to direct the fuel appropriately.

[0217] Although 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 catalysis inhibitors, biocides, static reducers, anti-icing agents, corrosion inhibitors) and any impurities. In addition to varying from airport to airport and from fuel supplier to fuel supplier, even for a given airport or fuel supplier, the fuel composition of available aviation fuel can vary from batch to batch.In addition, fuel tanks 50, 53 of aircraft 1 are generally not. not emptied before being refilled for a later flight, resulting in mixtures of different fuels in the tanks - a fuel of different composition effectively resulting from the mixing.

[0218] The inventors appreciated that, since different fuels can have different properties while still complying with standards, knowledge of the fuel(s) available for an aircraft 1 could allow for more efficient and tailored control of the propulsion system 2. For example, a fuel with a higher calorific value can be used to cool the engine more than a fuel with a lower calorific value, and a fuel with a higher calorific value can allow a lower fuel flow rate to be supplied to the combustion chamber for the same power output. Fuel knowledge can therefore be used as a tool to improve aircraft performance. In particular, the inventors appreciated that the programming of the Steerable Inlet Guide Vane (SIGGV) could be adjusted based on the fuel characteristics.

[0219] One or more fuel characteristics of a fuel arranged to be supplied to a gas turbine engine 10 of an aircraft 1 can therefore be obtained or otherwise determined and used to influence the control of the propulsion system 2; this can be described as the realization of a change in the operation of the propulsion system 2.

[0220] As used herein, the term "fuel characteristics" refers to intrinsic or inherent properties of the fuel such as fuel composition, and not to variable properties such as volume or temperature. Examples of fuel characteristics include one or more of: i. the percentage of sustainable aviation fuel (%SAF, by weight or by volume) in the fuel, or an indication that the fuel is a fossil fuel, for example fossil kerosene, or that the fuel is pure SAF; ii. the parameters of a hydrocarbon distribution of 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; • information on the percentage composition of some or all of the hydrocarbons present; iii. the presence or percentage of a particular element or species, such as: • the percentage of nitrogen-containing species in the fuel; • the presence or percentage of a tracer species or trace element / substance in the fuel (for example, a trace substance intrinsically present in the fuel which can vary from one fuel to another and therefore be used to identify a fuel, and / or a substance deliberately 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 when used in a gas turbine engine 10, such as: • the level of non-volatile particulate matter (nvPM) or CO2 emissions during combustion (a value can be provided for a specific combustion chamber operating under particular conditions in order to compare fuels fairly - a measured value can be adjusted accordingly based on the conditions and properties of the combustion chamber) • the level of fuel coking; v. one or more properties of the fuel itself, independent of its use in an engine 10 or its 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.

[0221] For example, the calorific value of a fuel may be selected as a fuel characteristic of interest. As used here, the term "calorific value" means the lower heat value (also called the lower heating value) of the fuel, unless otherwise specified. The lower heating 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).

[0222] The calorific values ​​(also called thermal values) 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; for example, based on the hydrocarbon distribution of the fuel and the calorific value of each type of constituent hydrocarbon (for which a standard value can be consulted). Alternatively, or in addition, 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.

[0223] A change of operation is a change in the current or intended operation of the propulsion system 2. In particular, changes to the programming of the Steerable Inlet Guide Vane can be made, based on the fuel characteristic(s) obtained. For example, a Steerable Inlet Guide Vane (SIGG) 246, as shown in [Fig. 4], can move in a direction and / or by a quantity determined based on the fuel characteristic(s). Alternatively, a SIGG can be held stationary in a condition / at a time when it would normally move, based on the fuel characteristic(s) different from those of a standard or previously used fuel.A change in operation may therefore, in some cases, be a decision not to perform a VIGV programming change that would normally be made under those circumstances (for example, a change in fuel flow or a change in aircraft speed). Examples of changes in operation thus include adjusting or canceling a VIGV positioning adjustment.

[0224] It will be understood that a change in VIGV geometry can generally be triggered by a change in the speed of the aircraft 1, a change in temperature at the inlet of a compressor 14, and / or a change in pressure across a compressor 14. The inventors appreciated that changes in VIGV geometry may also be appropriate when a fuel having different characteristics is used - as such, when a fuel is changed in flight (for an aircraft 1 with a plurality of different fuels on board) or between flights, a different VIGV programming may be appropriate even if all environmental and engine control factors other than the fuel are the same.

[0225] For example, for a given gravimetric fuel flow rate and shaft speed, the VIGVs can be opened more widely when using a fuel with a higher % S AF. Opening the VIGVs for a fuel with a higher % S AF or higher calorific value can achieve one or more of the following: improve efficiency, reduce T41, increase P30, and / or increase the overall pressure ratio through the compression system.

[0226] It will be understood that the geometry of VIGV / 1'opening angle can be measured directly, for example by using feedback from one or more angle control devices (e.g., actuator 242 described below), or can be deduced from side effects.

[0227] Changing the geometry of the VIGV changes the angle of airflow in the compressor 14. If one or more VIGVs 246 are not properly adjusted, the improper flow may result in compressor surge or stall unless corrective measures are taken (e.g., opening or closing a bleed valve, and / or making a further change in engine operation 10). Compressor stall is a local disturbance of the airflow in the compressor. Compressor surge is a stall that results in a complete disturbance of the airflow through the compressor 14. The severity of a stall ranges from a momentary and insignificant drop in power to a complete loss of compression in the case of surge, requiring adjustments to the fuel flow to recover normal operation.Monitoring pressures and flow rates makes it possible to detect when a compressor 14 is approaching a pumping point, and corrective action can then be taken (e.g., VIGV changes and / or purge valve changes).

[0228] A compressor 14 will only pump air stably up to a certain engine pressure ratio (the Engine Pressure Ratio (EPR) is the ratio of the turbine discharge pressure divided by the compressor inlet pressure); if the EPR is exceeded, the airflow will become unstable. This occurs at what is called the pump line on a compressor map. The engine 10 is designed to keep the compressor 14 running a short distance below the pump line, on an operating line of a compressor map. The distance between the two lines can be called the pump margin. A change in fuel characteristics can increase or decrease the operating pressure ratio, thus moving the operating line toward or away from the pump line. If the gap between the lines / pump margin decreases to zero, compressor stalling may result.

[0229] Modern compressors 14 are designed and controlled, usually by an electronic motor controller (ECC) 42, to avoid or limit stalling within the operating range of a motor.

[0230] Fig. 4 illustrates the airflow A, approaching a compressor 14, and more particularly the low-pressure compressor 14 of the gas turbine engine 10. The compressor 14 comprises a rotor having a plurality of blades 14a extending from a central region and arranged to work on the airflow passing through them.

[0231] In the implementation illustrated in [Fig.4], there is a plurality of VIGV 246 arranged in the working fluid flow path upstream of / at or near a compressor inlet 14. The VIGV 246 blade shown is only one of a plurality of VIGV 246s arranged around the fluid flow path in this example. The VIGV 246s are regularly spaced around the annular flow path in the example shown and can pivot to adjust the angle of the VIGVs relative to the fluid flow A. The VIGV arrangements may differ in other examples.

[0232] In the example shown in [Fig. 4], the plurality of VIGV 246 are coupled to an annular element 244, which allows the plurality of VIGV 246 to move in unison. An actuator 242 is functionally coupled to the annular element 244. The actuator 242 is controlled by the motor control system (EEC 42) and moves the annular element 244 the desired amount to effect a change in the position of the plurality of VIGV 246 relative to the fluid flow in the working fluid path. The actuator 242 may also include a position sensing function to provide feedback on the actual position of the VIGV 246. In an alternative example, a separate position sensor may be used to provide an output signal indicative of the actual position of the VIGV 246.It will be understood that different control and actuation arrangements can be used in different examples, for example with one or more VIGV 246 units that can be controlled independently.

[0233] A VIGV 240 programming management device is used to adjust the VIGV programming based on one or more fuel characteristics. One or more fuel characteristics are thus obtained for the fuel in order to perform the programming adjustment.

[0234] For a given fuel flow rate, fuel characteristics such as the fuel calorific value affect the temperature at the turbine inlet, and therefore the engine temperatures, pressures, and pressure / temperature ratios. Calorific value can thus be selected as one, or the, fuel characteristic on which VIGV programming changes are based.

[0235] In some examples, such as that shown in [Fig.6], the aircraft 1 may have only one fuel tank 50, and / or may have multiple fuel tanks 50, 53 which each contain the same fuel, and / or are fluidically connected, or are in fluidic communication with the gas turbine engine 10, so that only one type of fuel is supplied to the gas turbine engine 10 between refueling events - i.e. the fuel characteristics may remain constant throughout a flight and change only between flights.

[0236] In other examples, however, aircraft 1 may have a plurality of tanks of fluidly separated fuels 50, 53 containing fuels of different compositions, and the propulsion system 2 may include an adjustable fuel distribution system, allowing selection of the tank(s) 50, 53, and therefore which fuel / fuel mixture to use. In such implementations, the fuel characteristics may vary during a flight, with a specific fuel or fuel mixture being supplied to the gas turbine engine 10. The fuel characteristics for the multiple different fuels in each tank 50, 53 can therefore be determined, and / or the fuel characteristics of a fuel / fuel mixture currently supplied to the gas turbine engine 10 can be directly detected or otherwise determined.

[0237] Fuel characteristics, such as calorific values, can therefore be obtained in various 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 the fuel data, or a tracer substance (e.g. a dye) is identified and the fuel properties are consulted on the basis of this tracer; • data can be entered manually, or transmitted to aircraft 1 for storage; • a fuel sample can be extracted for ground analysis before takeoff; • Fuel properties can be deduced from measurements of propulsion system 2 activity during one or more aircraft operating periods, for example, engine start, 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.

[0238] 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, or by reference to data from a specific tracer detected in the fuel). The 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 compound or cycloparaffin content of the fuel can be determined based on measurements of the swelling of a component of sensor made from a sealing material such as a nitrile sealing material. Trace substances or species, naturally occurring in fuel or added to act as a tracer, can be used to determine fuel characteristics such as the percentage of sustainable aviation fuel in the fuel 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 the fuel's aromatic compound content, fuel oxygen content, and the fuel's thermal stability or coking level - for example, by measuring the accumulation of surface deposits on the piezoelectric crystal that will cause a change in vibrational mode. Various fuel characteristics can be determined by collecting performance parameters of the gas turbine engine 10 during an initial operating period (such as during takeoff), and possibly also during a second operating period (e.g., during cruise), and comparing these collected parameters to expected values ​​when using fuel with known properties. 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 condensation trail formation 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 hydrocarbon content and hydrogen-to-carbon ratio, can be determined by measuring the 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 aircraft 1, on the basis of measurements taken as the fuel is burned - for example by using the fuel flow rate and shaft speed or the temperature change across the combustion chamber 16. Various fuel characteristics can be determined by performing a change of operation arranged to affect the operation of the gas turbine engine 10, by detecting a response to the change of operation; and by determining the fuel characteristic(s) of the fuel based on the response to the change of operation. • 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 a second fuel, and by determining the fuel characteristic(s) of the second fuel based on a change in the relationship between T30 and one of T40 and T41 (the relationship indicating the temperature rise across the combustion chamber 16). The characteristics can be determined as relative values ​​with respect to the first fuel, or as absolute values, for example by reference to known values ​​for the first fuel.

[0239] In examples where a fuel cannot be changed in flight, the VIGV 240 scheduling management device can be provided with a list of one or more fuel characteristics, which list is then used throughout the flight / until the next refueling event. The fuel characteristic(s) are thus obtained only once per flight or refueling event and used multiple times throughout the flight, each time a VIGV 246 movement is planned or contemplated.

[0240] In examples where a fuel or fuel mixture can be changed in flight, the fuel characteristic(s) of the fuel supplied to the combustion chamber 16 can change during flight as the fuel or fuel mixture is changed, so that values ​​can be obtained multiple times during a flight. For example, the VIGV 240 programming management device can obtain values ​​for the fuel characteristics (i) at regular intervals (possibly with the frequency varying according to the stage of flight, for example, less frequently during cruise than during climb); (ii) each time the fuel or fuel mixture supplied to the gas turbine engine 10 is changed; and / or (iii) before each (potential) VIGV programming change.

[0241] The VIGV 240 programming management device can obtain data on a percentage mixture of one or more different fuels supplied to the gas turbine engine 10 at a certain time, consult fuel characteristic data for the / each fuel in a data store, and determine / calculate the fuel characteristics for the fuel / mixture accordingly. In some examples, no in-flight detection or analysis can be performed, and We can instead rely on pre-supplied data. In other examples, physical and / or chemical detection (either of the fuel characteristic(s) directly, or of one or more fuel or engine properties from which the fuel characteristic(s) can / can be derived) can be used instead of, or in addition to, retrieving data from storage.

[0242] The programming management device of VIGV 240 is therefore arranged to obtain one or more characteristics of the fuel currently supplied to the gas turbine engine 10 in any appropriate manner.

[0243] Once one or more fuel characteristics have been determined for a fuel currently supplied to the gas turbine engine 10, the propulsion system control 2, and in particular the VIGV programming, can be adjusted on the basis of the determined fuel characteristic(s). It will be understood that, for many current aircraft 1, changes to the VIGV programming may only be applicable to geared gas turbine engines 10.

[0244] For example, for a 2% increase in the calorific value of a fuel supplied to the gas turbine engine 10, the VIGVs can be opened at takeoff by about 2% of their range (assuming a full range of motion / rotation of 40°). For example, for a given aircraft 1 with a usual VIGV angle for Jet A, the VIGVs can be opened beyond this usual angle by 5% of their range (i.e., displaced by 2°) if a fuel having a calorific value 5% higher than that of Jet A is used. This change in VIGV programming can facilitate the maintenance of a more constant turbine gas temperature (e.g., T41). A corresponding change can be made in cruise, although the magnitude of the position change is likely to be smaller.It will be understood that VIGV programming changes can be adapted to a particular aircraft 1, and / or to a particular part of the flight envelope (e.g., takeoff or cruise), so as to achieve a certain turbine gas temperature (e.g., T41), or a certain temperature rise through the combustion chamber 16 (e.g., the relationship between T30-T41).

[0245] As a further example, for a 30% increase in heat capacity, the VIGV 246 engines can be opened an additional 0.5% at takeoff, up to a limit of 5% of their full range. This can be scaled linearly for a smaller (or larger) change in heat capacity. A corresponding change can be achieved in cruise, although the magnitude of the change is likely to be smaller. Similarly, a 30% decrease in heat capacity can cause the VIGV 246 engines to be closed by 0.5% at takeoff, up to a limit of 5% of their full range.

[0246] Additional data can be used in conjunction with the determined fuel characteristics to adjust the control of the VIGV 246. For example, the described approach may include receiving operating parameter data such as aircraft speed, air and / or fuel flow rate, temperature at the inlet of a compressor 14, and / or pressure across a compressor 14, fuel temperature data, and / or environmental parameters such as altitude. This received data (e.g., operating and / or environmental parameters) can be used to make or influence changes to the VIGV programming. For example, if the fuel temperature were higher at the inlet of the combustion chamber 16, for every 50-degree increase in fuel temperature at takeoff, the VIGV 246 can be opened by 1%.

[0247] A propulsion system 2 for an aircraft 1 can therefore include one or more steerable inlet guide vanes -VIGV- 246 through / in front of which the airflow passes into the compressor 14; and a VIGV programming management device 240 arranged to obtain one or more characteristics of the fuel supplied to the gas turbine engine 10; and effect a programming change of the or more VIGV 246 on the basis of the or more characteristics of the fuel obtained.

[0248] The VIGV 240 programming management device can determine a desired change in VIGV programming based on the one or more fuel characteristics obtained and control an actuator 242 so as to move the one or more VIGV 246 accordingly.

[0249] In the implementation shown in [Fig.4], a separate VIGV programming management device 240 is provided for each gas turbine engine 10. In other implementations, a single VIGV programming management device 240 may be provided and may control the VIGV programming for both (or all) engines 10.

[0250] The VIGV programming management device 240 of the illustrated example also includes a receiver 241 arranged to receive data relating to a fuel composition and / or requests for VIGV programming changes. The determination of a desired VIGV programming change can therefore be carried out by the VIGV programming management device 240 itself, or the VIGV programming management device 240 can implement a change determined by another entity, depending on the implementation.

[0251] A fuel composition monitoring device 202 can be used to record and store fuel composition data, and optionally also to receive sensor data (and possibly other data) and to calculate fuel characteristics based on this data. The VIGV 240 programming management device can be provided as part of the same entity, or can obtain data from the fuel composition monitoring device 202.

[0252] The fuel composition monitoring device 202 of the described example includes a memory 202a (which may also be called computer storage) arranged to store current fuel characteristic data, 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 can be time-stamped and / or date-added to the memory. A log of fuel characteristic data over time can thus be assembled.

[0253] The fuel composition monitoring device 202 of the illustrated example also includes a receiver 202b arranged to receive data from which fuel characteristics can be calculated, and / or the fuel characteristics themselves, and / or requests for fuel composition information. The fuel composition monitoring device 202 of the illustrated example is part of, or in communication with, an electronic engine controller (EEC) 42. The EEC 42 can be arranged to issue propulsion system control commands based on the calculated fuel characteristics. It will be understood that an EEC 42 may be provided for each gas turbine engine 10 of the aircraft 1, or that a single EEC 42 may control both or all of the engines 10. In addition, the role played by the EEC for the fuel composition monitoring device 202 may represent only a small part of the functionality of the EEC.Indeed, the fuel composition monitoring device 202 may be supplied by the EEC, or may include a separate EEC module from the engine's EEC 42 in various implementations. In alternative examples, the fuel composition monitoring device 202 may not include any engine control functionality, and may instead simply provide fuel composition data on demand, for appropriate use by another system. Optionally, the fuel composition monitoring device 202 may provide a proposed change to the engine control functionality for approval by a pilot (or other authority); the pilot may then directly implement the proposed change, or approve or reject the automatic implementation of the proposed change.

[0254] The propulsion system 2 may therefore include an electronic motor controller 42 arranged to issue propulsion system control commands on the basis The fuel characteristics are determined based on data provided by the fuel composition monitoring device 202 and / or the VIGV programming management device 240, and possibly other data. The VIGV programming management device 240 in the example shown may be part of, or communicate with, the engine electronic controller (EEC) 42, which is configured to issue propulsion system control commands based on the fuel characteristics. It will be understood that the role of the EEC 42 for the VIGV programming management device 240 may represent only a small part of the EEC's functionality. Indeed, the VIGV programming management device 240 may be provided by the EEC 42, or it may comprise a separate EEC module from the engine's EEC 42 in various implementations.In alternative examples, the VIGV 240 programming management device may not include any engine control functionality, but may instead provide VIGV programming data on demand for appropriate use by another system. The fuel composition monitoring device 202 and / or the VIGV 240 programming management device may be provided as a separate unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into other aircraft control systems such as EEC 42. The fuel composition monitoring means may be provided as part of the same unit or assembly as the engine control functionality.

[0255] The EEC 42, which can also be considered a propulsion system control device, can make changes to the propulsion system 2, and in particular to the VIGV programming, directly, or can provide a notification to the pilot (or other authority) recommending the change, for approval. In some examples, the same propulsion system control device 42 can automatically perform some changes and request others, depending on the nature of the change. In some examples, the same implementation can include 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 to engine flow rates that do not affect engine power output and would not be noticed by a pilot—can be made automatically, whereas any change that the pilot would notice can be notified to the pilot (i.e., a notification appears 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 implementations where a notification or suggestion is provided to a pilot, this may... be provided on an aircraft cockpit display and / or as an audible alarm, and / or sent to a separate device such as a handheld tablet or other computing device.

[0256] A method 3010 for controlling a propulsion system 2 of an aircraft 1 can therefore be implemented, the propulsion system 2 comprising a gas turbine engine 10 with one or more VIGV 246 at or near the inlet of a compressor 14 of the gas turbine engine 10.

[0257] The process 3010 includes obtaining 3012 one or more characteristics of the fuel supplied to the gas turbine engine 10. Obtaining 3012 can be carried out by retrieving data from storage and / or by physically and / or chemically detecting one or more fuel properties. The obtaining step 3012 can be carried out only once, for example during refueling or at the start of a flight. Particularly in examples where a fuel or fuel mixture can be changed in flight, the obtaining step 3012 can be carried out repeatedly during a flight.

[0258] The method 3010 includes the realization 3014 of a change in the programming of the VIGV or more 246 based on the characteristic or characteristics obtained from the fuel, for example by moving a VIGV by a certain amount (for example a rotation by a certain angle), in a certain direction.

[0259] In implementations with variable fuel in flight, the data acquisition step 3012 and the data change implementation step 3014 can be repeated together each time a change in VIGV position is considered, or the data acquisition step 3012 can be performed at intervals. In implementations with a single constant fuel in flight, the data acquisition step 3012 can be performed only once and the data change implementation step 3014 can be performed several times during a flight, using the same data obtained. Alternatively, the data acquisition step 3012 can again be performed at intervals, for example, for verification.

[0260] As described above, the inventors appreciated that knowledge of the fuel(s) available for an aircraft 1 could allow for more efficient and tailored control of the propulsion system 2—as with the VIGV programming control described herein. In some cases, the fuel characteristics can be provided to the aircraft 1 by a third party, for example, by a supplier during refueling. However, in other cases, prior knowledge of the fuel characteristics may not be available. One or more fuel characteristics of a fuel arranged to be supplied to a gas turbine engine 10 of an aircraft 1 can therefore be determined on board the aircraft 1 and possibly subsequently used to influence the control of the propulsion system 2.

[0261] In the examples described below, the aircraft propulsion system 2 is used to perform an "experiment" to determine, or to provide data useful for determining, one or more fuel characteristics. This performance of an "experiment" includes carrying out a change in the operation of the propulsion system 2 and determining the effect of this change in operation—one or more fuel characteristics can then be determined from the response to the known change in operation. The fuel characteristics may include one or more of those listed above.

[0262] More specifically, a change of operation is effected, the change of operation being effected by a controllable component of the propulsion system 2. The change of operation is selected to affect the operation of the gas turbine engine 10 in a manner dependent on at least one fuel characteristic.

[0263] A change of operation is a change in the current or intended operation of the propulsion system 2. For example, a steerable inlet guide vane (VIGV) 246 may move, and a response to this movement is detected. Alternatively, a VIGV may be held stationary in a condition / at a time when it would normally move, and a response to this change relative to the standard operating procedure may be monitored. A change of operation may therefore, in some cases, be a decision not to perform a change of operation that would normally be performed under those circumstances.It will be understood that this can be considered the inverse of the 3012, 3014 approach described above - rather than obtaining one or more fuel characteristics and changing the VIGV programming based on these fuel characteristics to obtain a desired response, a change in the VIGV programming is made and one or more fuel characteristics are deduced or determined from the response to this programming change.

[0264] For example, the VIGV 246 can move in such a way as to maintain a constant temperature T41 or relationship between T30-T41 during fuel change (e.g., T41 less T30 or T40 less T30, indicating a temperature rise in the combustion chamber); the movement required to maintain the temperature or relationship between temperatures constant can then be used to identify a change in calorific value between the initial fuel (before the change of fuel supplied to the gas turbine engine 10) and the new fuel.

[0265] Assuming that the fuel mass flow rate is kept constant during the fuel change, an increase in the temperature rise across the combustion chamber 16 (T40-T30) is likely to be observed during a passage to a fuel with a higher calorific value if no change is made to the VIGV programming. If it is decided not to change the VIGV programming by changing the fuel / observing that the temperature rise begins to increase, the change in temperature rise across the combustion chamber 16 can be used to calculate the change in the fuel's calorific value. For current SAF and SAF blends, a change in temperature rise of at least 2% or 3% can be observed compared to kerosene, which can correspond to a change of more than 30°C, or more than 50°C.

[0266] If the low-pressure shaft speed / thrust is kept constant instead of the fuel mass flow rate, an increase in T41 may still be observed due to the higher calorific value of the new fuel if no change is made to the VIGV programming, and the magnitude of this change can be used to deduce the change in calorific value. A change of approximately 3°C can be observed for every 3% change in fuel calorific value.

[0267] As described above, a compressor 14 will only pump air stably up to a certain engine pressure ratio (the Engine Pressure Ratio (EPR) is the ratio of the turbine discharge pressure (P42) divided by the compressor inlet pressure (P26)); if the EPR is exceeded, the airflow will become unstable. This occurs at what is called the pump line on a compressor map. The engine is designed to keep the compressor running a short distance below the pump line, on an operating line of a compressor map. The distance between the two lines can be called the pump margin. A change in fuel characteristics can increase or decrease the operating pressure ratio, thus moving the operating line toward or away from the pump line.If the gap between the lines / pumping margin decreases to zero, a compressor stall may result.

[0268] Modern compressors 14 are designed and controlled, usually by EEC 42, to avoid or limit stalling within an engine's operating range. While compressor surge should generally be completely avoided, the precise point at which a small stall occurs for a given fuel flow rate can be used to deduce the fuel characteristics. The compressor 14 will then return to a normal flow rate once the engine pressure ratio is reduced to a level at which the compressor can maintain a steady airflow.

[0269] For example, for a given fuel flow rate, the fuel's calorific value affects the temperature at the turbine inlet, and therefore the engine's pressure / temperature ratios. Monitoring the proximity of compressor stall 14 after changing the geometry of VIGV, or after changing the fuel and not changing the geometry of VIGV, can therefore allow us to determine or deduce a calorific value or another parameter of the fuel.

[0270] While airflow patterns can be measured in some implementations, VIGV angles and side effects such as temperature and pressure changes can be more easily measured directly. For example, as well as changes in the T30-T41 relationship, opening the VIGVs 246 often results in a higher P30 and an increase in the overall pressure ratio across the compression system. Furthermore, VIGV position information can be directly available from one or more actuators 242.

[0271] Other examples of changes in operation, apart from changes in VIGV programming, may include adjusting or canceling an adjustment of one or more of the following: • fuel composition (for example, variation in a % mixture of fuels from two different sources / tanks 50, 53); • fuel temperature (e.g., the temperature of the fuel entering the combustion chamber 16) or one or more other thermal management characteristics; • the engine's thrust; • fuel flow rate; • the fuel pump spillage rate; and • the injection of water into the combustion chamber 16.

[0272] For example, if a fuel change is carried out while the gas turbine 10 is maintained running at a fixed speed / thrust and the fuel mass flow rate has decreased but not the volumetric flow rate, then it can be deduced that the new fuel has a lower density, and the density can be calculated accordingly. It will be understood that, for many current flow sensors, a change in flow rate can be more accurate than an absolute value, thus allowing the density to be calculated more accurately during the fuel change, by reference to values ​​for the first fuel, than would be possible using sensor flow values ​​for a single fuel.

[0273] By way of further example, if the airflow and / or oil flow to one or more air-oil heat exchangers 118 is / are reduced during the fuel change and no pressure increase (or a lower pressure increase than expected for the original fuel) is observed through all or part of the fuel circuit 3 and / or if no change in fuel temperature (or a smaller change in fuel temperature than expected for the original fuel) is observed, it can be inferred that the The new fuel has better calorific value and / or thermal stability (the absence of pressure increase indicating the absence of carbon deposit formation). (Fuel circuit 3 includes the fuel route between tanks 50, 53 and the engine(s) 10, including all pipelines and components along this route.) It will be understood that reducing the airflow to the air-oil heat exchanger 118 (which may be called an air cooler) would result in less oil cooling and, consequently, less heat removal from the engine 10, and therefore a hotter engine 10 and more heat in the fuel, and that reducing the oil flow to the air-oil heat exchanger 118 may cause more hot oil to be directed to a fuel-oil heat exchanger (not shown), thus directly adding heat to the fuel.

[0274] By way of further example, in a gas turbine engine 10 comprising a combustion chamber 16 with multiple different combustion modes, a change in nvPM generation can be monitored when a change is made between the combustion modes - the observed change in nvPM generation can be used to determine one or more fuel characteristics, for example the percentage of SAF or the nvPM generation potential itself.

[0275] Multiple operating changes can be made simultaneously, or successively, and the behavior of the propulsion system 2 can be monitored over a period of time, gathering data to determine the fuel characteristic or characteristics of interest.

[0276] In some examples, the aircraft 1 may have only one fuel tank 50, and / or may have multiple fuel tanks 50, 53 which each contain the same fuel, and / or are fluidically connected, or are in fluidic communication with the gas turbine engine 10, so that only one type of fuel is supplied to the gas turbine engine 10 between refueling events - i.e. the fuel characteristics may remain constant throughout a flight.

[0277] In other examples, the aircraft 1 may have a plurality of fuel tanks 50, 53 containing fuels of different compositions, and the propulsion system 2 may include an adjustable fuel distribution system, allowing selection of the tank(s) 50, 53, and therefore which fuel / fuel mixture to use. In such examples, fuel characteristics may vary during a flight, and a specific fuel or fuel mixture may be selected to improve operation at certain stages of flight or under certain external conditions. In such examples, the same change in operation may be made at several different times, an active fuel management system 214 being arranged to change the fuel, or the mixture of fuels, between these moments. Fuel characteristics for the multiple different fuels on board can therefore be determined.

[0278] For example, in implementations in which the fuel temperature at the inlet of the combustion chamber 16 is changed, a response to this change in operation may be or include (i) a change in the power output of the gas turbine engine 10; or (ii) a change in the degradation or coking of the fuel.

[0279] Once one or more fuel characteristics have been determined for a fuel currently supplied to the gas turbine engine 10, the control of the propulsion system 2 can be adjusted on the basis of the determined fuel characteristics.

[0280] Additional data can be used in conjunction with the determined fuel characteristics to adjust the control of the propulsion system 2. For example, the method may include receiving data on the current conditions around the aircraft 1 (either from a supplier, such as a third-party weather monitoring agency, or from onboard sensors). This received data (e.g., weather data, temperature, humidity, contrail presence, etc.) can be used to make or influence changes to the propulsion system control. Instead of, or in addition to, using "real-time" or near-real-time weather data, forecast weather data for the aircraft's route can also be used to estimate current conditions.

[0281] By way of further example, in implementations in which the propulsion system 2 includes a plurality of unconnected fluidly coupled fuel tanks 50, 53, the achievement of a change of operation may include changing or consist of changing the tank 50, 53 from which fuel is drawn, or changing the percentage of fuel drawn from a particular tank, thereby changing the fuel composition.

[0282] The response to a change in fuel composition may consist of or include one or more of the following examples:

[0283] (i) a change in the output power of the gas turbine engine 10;

[0284] (ii) a change in the degradation or coking of the fuel;

[0285] (iii) a change in the formation of condensation trails (contrails can be detected visually and / or by an infrared sensor, or can be deduced from measurements of temperature, pressure and humidity, among other variables, for example);

[0286] (iv) a change in the Engine Pressure Ratio;

[0287] (v) a change in the relationship between a temperature at the compressor outlet - T30 - and a temperature at the turbine rotor inlet - T41;

[0288] (vi) a change in the relationship between a total pressure at the compressor outlet - P30 - and a total pressure at the turbine rotor inlet - P41.

[0289] In the examples described, a turbine 17 of the engine 10 comprises a rotor having a leading edge and a trailing edge. A temperature at the turbine rotor inlet -T41- is defined as an average airflow temperature at the leading edge of the turbine rotor 17 under cruising conditions. Similarly, a pressure at the turbine rotor inlet -P41- is defined as the total airflow pressure at the leading edge of the turbine rotor 17 under cruising conditions.

[0290] The engine 10 also includes a compressor 15 having an outlet, and a temperature at the compressor outlet - T30 - is defined as an average airflow temperature at the outlet of the compressor 15 under cruising conditions. Similarly, a pressure at the compressor outlet - P30 - is defined as the total airflow pressure at the outlet of the compressor 15 under cruising conditions. In some examples, the gas turbine engine 10 includes multiple compressors; the temperature or pressure at the compressor outlet can be defined as the temperature or pressure at the outlet of the compressor 15 at the highest pressure. The compressor 15 may include one or more rotors, each having a leading edge and a trailing edge; the temperature or pressure at the compressor outlet can be defined as the temperature or pressure at the axial position of the trailing edge of the rearmost rotor of the compressor.

[0291] Between station 40 (the combustion chamber outlet) and station 41 (the inlet of the high-pressure turbine 17) there is generally provided a set of turbine distributor blades which can move to modify the flow in the rotating turbine 17; these are often described as steerable inlet guide vanes - VIGV 246 - as described above.

[0292] Once one or more fuel characteristics have been determined for a fuel currently supplied to the gas turbine engine, the control of the propulsion system 2 can be adjusted on the basis of the determined fuel characteristics.

[0293] In addition or alternatively, a planned flight profile can be changed on the basis of the one or more determined fuel characteristics.

[0294] As used herein, the term "flight profile" refers to the operational characteristics (e.g., height / altitude, power setting, flight path angle, aerodynamic speed, and the like) of an aircraft 1 when flying along a flight path, as well as the flight path / route (route) itself. Changes in route are therefore included in the term "flight profile" as used here.

[0295] Additional data may be used in conjunction with the determined fuel characteristics to adjust the control of propulsion system 2 and / or flight profile changes, as described above with respect to the control of propulsion system 2.

[0296] Once the fuel characteristic or characteristics of the resulting fuel in the fuel tank 50, 53 after refueling have been determined, the propulsion system 2 can be controlled on the basis of the calculated fuel characteristics.

[0297] For example: • An operating parameter of an aircraft thermal management system (e.g., a fuel-oil heat exchanger or an air-oil heat exchanger 118) can be changed, or the temperature of fuel supplied to the combustion chamber 16 of the engine 10 can be changed. • When more than one fuel is stored on board an aircraft 1, the selection of the fuel to be used for a given operation (e.g., for ground operations as opposed to flight, for low-temperature starting, or for operations with different thrust requirements) or at a given time during flight can be made based on fuel characteristics such as % SAF, nvPM generation potential, viscosity, and calorific value. A fuel distribution system can therefore be appropriately controlled based on these fuel characteristics. • One or more aircraft 1 control surfaces can be adjusted to change route and / or altitude based on knowledge of fuel. • The spillage percentage of a fuel pump (i.e., the proportion of pumped fuel recirculated instead of being delivered to the combustion chamber) can be changed, for example, based on the fuel's SAF percentage. The pump and / or one or more valves can therefore be appropriately controlled based on the fuel characteristics. • Changes to the programming of the steerable inlet guide vanes (VIGV 246) can be made based on fuel characteristics. The VIGV 246 vanes can therefore be moved, or a VIGV movement can be canceled, as appropriate, based on fuel characteristics.

[0298] A propulsion system 2 for an aircraft 1 may therefore include a fuel composition tracking device 202 arranged to record and store fuel composition data, and optionally also to receive data of a change of operation and measurement data relating to a response to the change of operation and to calculate one or more fuel characteristics on the basis of this data (and optionally also on the basis of other data, such as measurement data relating to responses to one or more other changes of operation, or reference tables).

[0299] The fuel composition monitoring device 202 may be provided as a separate fuel composition monitoring unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems.

[0300] Data from the fuel composition monitoring device 202 can be used to adjust the control of the propulsion system 2, based on the fuel characteristic or characteristics.

[0301] In the example shown, two sensors 204 are provided, each arranged to physically and / or chemically detect one or more performance characteristics of the gas turbine engine. In different implementations, different numbers and / or types of sensors may be provided. For example, one or more pressure and / or temperature sensors 204 may be provided, a fuel flow sensor may be provided, and / or one or more chemical sensors may be provided, for example, to detect exhaust characteristics or fuel components. The sensors 204 and the fuel composition monitoring device 202 may be described together as a fuel composition monitoring system 203, as shown in [Fig. 8]. In some implementations, pre-existing sensors may be used so that the implementation of the method 2090 described below may not require any hardware changes.In other implementations, one or more additional sensors may be added to the propulsion system 2.

[0302] The fuel composition monitoring system 203 includes a fuel composition monitoring device 202, or another fuel composition determination module 210. The fuel composition monitoring device 202 of the described example includes a memory 202a arranged to store current fuel characteristic data, and processing circuits 202c arranged to calculate updated values ​​for the one or more fuel characteristics of the fuel burned in the engine 10. The calculated values ​​can then replace the fuel characteristic data previously stored in the memory, and / or can be time-stamped and / or date-added to the memory. A log of Fuel characteristic data over time can therefore be compiled. In other implementations, a log may not be maintained, and indeed, instantaneous ordering decisions can be made without storing fuel composition data for an extended period. In such implementations, the term fuel composition determination module 210 may be preferred over a fuel composition tracking device 202, since past data may not be tracked—the terms may otherwise be used interchangeably.

[0303] In the implementation shown in [Fig.6], a separate fuel composition determination module 210 is provided for each gas turbine engine 10. In other implementations, only one fuel composition determination module 210 may be provided.

[0304] The fuel composition monitoring device 202, 210 of the example shown also includes a receiver 202b arranged to receive data relating to a fuel composition and / or requests for fuel composition information.

[0305] The propulsion system 2 may include an electronic engine controller 42 arranged to issue propulsion system control commands based on determined fuel characteristics, based on data provided by the fuel composition monitoring device 202 and possibly other data. The fuel composition monitoring device 202 of the illustrated example may be part of, or in communication with, the electronic engine controller (EEC) 42, and the EEC 42 may be arranged to issue propulsion system control commands based on the fuel characteristics. It will be understood that an EEC 42 may be provided for each gas turbine engine 10 of the aircraft 1, and / or that the role played by the EEC 42 in or for the fuel composition monitoring device 202 may represent only a small part of the functionality of the EEC.Indeed, the fuel composition monitoring device 202 can be provided by the EEC 42, or can comprise a separate EEC module from the engine's EEC 42 in various implementations. In alternative examples, the fuel composition monitoring device 202 may not include any engine control functionality, and may instead simply provide fuel composition data on demand, for appropriate use by another system. The fuel composition monitoring device 202 may be provided as a separate propulsion system control unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into other aircraft control systems. The fuel composition monitoring means may be provided as part of the same unit or the same [unclear]. together with the motor control functionality, or separately.

[0306] The EEC 42, which can also be considered a propulsion system control device, can make changes to the propulsion system 2 directly, or can provide a notification to the pilot recommending the change for approval, as discussed above. In some examples, the same propulsion system control device 42 can automatically perform some changes and request others, depending on the nature of the change, as discussed above.

[0307] The propulsion system control device 42 may also provide recommendations regarding flight profile changes. Alternatively, or in addition, the propulsion system 2 may further include a flight profile adjustment device arranged to change a planned flight profile based on the fuel's fuel characteristics and possibly other data. The flight profile adjustment device may be provided as a separate propulsion system control unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems. The fuel composition monitoring means may be provided as part of the same unit or assembly.

[0308] A method 2090 for determining one or more fuel characteristics of a fuel supplied to a gas turbine engine 10 of an aircraft 1 can therefore be implemented, the gas turbine engine 10 being part of a propulsion system 2.

[0309] The process 2090 includes the implementation 2092 of a change in operation, the change in operation being caused by a controllable component of the propulsion system 2 and arranged to have a measurable effect on the operation of the gas turbine engine 10. The change in operation is any appropriate change in the operation of the propulsion system that will have an effect on the operation of the gas turbine engine 10, and may be or include the movement of a component of the propulsion system 2 (for example, the movement of a VIGV, the change in pump speed, the diversion of fuel and / or the opening of a purge valve), or may be or include the non-movement of a component of the propulsion system 2 in a situation where, according to normal operating procedures, it would normally move.The change in operation may be temporary and may be reversed as soon as sufficient time has elapsed for any effect on the operation of the gas turbine engine 10 to be detected (noting that a time interval may be left to allow any transient effect to dissipate in some cases, as described in more detail below).

[0310] The method 2090 further includes the detection 2094 of a response to a change in operation - for example, a change in one or more pressures, tem Temperatures, shaft speeds, and / or ratios such as the engine pressure ratio. Alternatively, or in addition, the change may be a change in condensation drag formation, coking, or any other appropriate engine parameter. The time response may be evaluated instead of, or in addition to, examining values ​​at defined times before and after the change.

[0311] The process 2090 further includes the determination 2096 of the one or more fuel characteristics of the fuel burned by the gas turbine engine 10 on the basis of the response to the change in operation.

[0312] In some implementations, the process 2090 may further include making one or more changes to the aircraft operation and / or the planned flight profile after the determination 2096 is made, based on the determined fuel characteristic(s), for example, to improve engine efficiency or reduce the impact on the climate (e.g., by adjusting the formation of condensation trails). In other implementations, knowledge of the fuel characteristics may not be used to change the aircraft operation, but may be used to influence refueling choices and / or to verify that the fuel data provided for a fuel are correct.In the event of a significant discrepancy between the determined and expected fuel characteristics, aircraft 1 may be returned to a refueling station for fuel verification, and / or further checks may be carried out. EEC 42 can be configured to provide a warning / alert to a pilot in such scenarios. In some implementations, the "experiment" can therefore be performed very early in aircraft operation—for example, during engine warm-up and / or other pre-taxi operations, or during the initial stages of taxiing—to facilitate a return to a refueling station if necessary.

[0313] The change of operation carried out in step 2092 may temporarily have an adverse (usually minor) effect on engine operation; for example, by decreasing efficiency or pushing the propulsion system 2 closer to the limits of its operating envelope. Such a temporary adverse effect on engine operation may be acceptable because of the improvements in engine performance that can subsequently be made once the fuel characteristics are known, by optimizing engine performance for the fuel type. In some implementations, the change of operation carried out in step 2092 may be performed while the engine 10 is idling with the aircraft 1 on the ground, so that flight operation is never adversely affected. In implementations with multiple fuel sources, the fuel or mixture supplied to the engine 10 may be changed while idling. to enable the determination and storage of one or more fuel characteristics of each stored fuel for later reference.

[0314] In implementations in which a fuel composition monitoring device 202 as described above is used to carry out process 2090, the fuel composition monitoring device 202 may be arranged to:

[0315] receive information concerning a change of operation, the change of operation being effected by a controllable component of the propulsion system 2 and arranged to affect the operation of the gas turbine engine 10;

[0316] receive data corresponding to a response to the change in operation; and

[0317] determine one or more fuel characteristics of the fuel arranged to be supplied to the gas turbine engine 10 on the basis of the response to change in operation, as determined from the data received.

[0318] In the examples described below, one or more temperatures and / or pressures inside the gas turbine engine 10 (and possibly a relationship between temperatures and / or pressures at different points inside the gas turbine engine 10) are used to determine or provide useful data for determining one or more fuel characteristics of the fuel currently being burned in the engine 10.

[0319] In particular, in examples using one or more temperatures, each temperature or the relationship between temperatures is recorded for a first fuel, and then recorded again after a change of fuel. A difference in fuel characteristics, for example, an increased calorific value, can thus be determined from a difference in temperature(s) or a relationship between temperatures. Instead of "performing an experiment" for a single fuel currently being burned, the change of fuel is the difference, and a response to the change of fuel is used to determine one or more fuel characteristics.

[0320] For example, T41, or a relationship between T30 and T41, can change depending on the % SAF of a fuel if the automatic VIGV adjustment (for example, to maintain T41 or the relationship between temperatures constant) is canceled or delayed. A change of approximately 5°C in T41 can occur, for example, if there is a change between kerosene and a currently used SAF. It will be understood that VIGV programming may traditionally be based on maintaining a constant level of one or more of T40, T41, T30, or the relationship between T30 and T41, and that allowing the temperature to change and observing by how much, rather than automatically shifting the VIGVs, can allow the fuel characteristics to be inferred.

[0321] Changes in temperature(s) or in the relationship between temperatures can be used to identify relative fuel characteristics, rather than absolute values—for example, an 8% increase in calorific value compared to the previous or reference fuel—in some examples. In other examples, absolute values ​​can be calculated, possibly by reference to data that may include absolute values ​​for the previous or reference fuel.

[0322] One or more pressures may also change - in some cases, both pressures and temperatures may be monitored, and a detected change in one is used to verify a detected change in the other.

[0323] In additional or alternative examples using pressures, one or more pressures and / or a relationship between pressures is / are noted for a first fuel, and then noted again after a change of fuel. A difference in fuel characteristics, for example, an increased calorific value, can thus be determined from a difference in pressure(s) or a relationship between pressures. As with temperature changes, changes in pressure(s) can be used to identify relative fuel characteristics, rather than absolute values—for example, an 8% increase in calorific value compared to the previous or reference fuel—in some examples. In other examples, absolute values ​​can be calculated, possibly by reference to data for the previous or reference fuel.

[0324] In various examples, pressures and temperatures are detected, measured, calculated or otherwise deduced, and both can be used to determine fuel characteristics.

[0325] The propulsion system 2 may include one or more steerable inlet guide vanes - VIGV 246 - and a fuel pump. No change in the position of the VIGV 246 and / or the fuel flow rate may be made during fuel change, at least until after the collection of updated temperature and / or pressure data, so as to allow monitoring of any temperature and / or pressure changes with minimal interference / minimal variation in engine control beyond the fuel type.

[0326] Multiple relationships between temperatures, between multiple gas turbine engine temperatures, can be used in some examples. In additional or alternative examples, multiple relationships between pressures, between multiple gas turbine engine pressures, can be used.

[0327] In the examples described, a combustion unit 16, for example being or comprising a combustion chamber 16, burns the fuel in the turbine engine The combustion chamber 16 has an outlet, and the temperature at the combustion chamber outlet – T40 – is defined as the average airflow temperature at the combustion chamber outlet under cruising conditions. Similarly, the pressure at the combustion chamber outlet – P40 – is defined as the total airflow pressure at the combustion chamber outlet under cruising conditions. The airflow from the combustion chamber 16 then enters a turbine 17.

[0328] In the examples described, a turbine 17 of the engine 10 comprises a rotor having a leading edge and a trailing edge. A temperature at the turbine rotor inlet -T41- is defined as an average airflow temperature at the leading edge of the turbine rotor 17 under cruising conditions. Similarly, a pressure at the turbine rotor inlet -P41- is defined as the total airflow pressure at the leading edge of the turbine rotor 17 under cruising conditions.

[0329] The engine also includes a compressor 15 having one outlet, and a temperature at the compressor outlet - T30 - is defined as an average airflow temperature at the outlet of the compressor 15 under cruising conditions. Similarly, a pressure at the compressor outlet - P30 - is defined as the total airflow pressure at the outlet of the compressor 15 under cruising conditions. In some examples, the gas turbine engine 10 includes multiple compressors 14, 15; the temperature or pressure at the compressor outlet can be defined as the temperature or pressure at the outlet of the compressor 15 at the highest pressure.The compressor 15 may comprise one or more rotors, each having a leading edge and a trailing edge; the temperature or pressure at the outlet of the compressor may be defined as the temperature or pressure at the axial position of the trailing edge of the rearmost rotor of the compressor.

[0330] One or more of the listed temperatures and / or pressures are used to determine one or more fuel characteristics. A change in the relationship between T41 and T30, and / or between P41 and P30, can be used to determine the fuel characteristic(s). T40 or P40 may be used in addition to or instead of T41 or P41 in certain examples.

[0331] In various implementations, cooling air at temperatures T30 can be introduced through a turbine distributor vane at the outlet of the combustion chamber 16, between stations T40 and T41. In some implementations, particularly those where the amount of cooling air added varies, T40 can be selected instead of T41 to avoid any variability in T41 due to the amount of cooling air influencing changes in the relationship / temperature.

[0332] As mentioned above, T30, T41, P30 and P41 and all other numbered pressures and temperatures listed herein are defined using the station numbering listed in SAE AS755, in particular: • P30 = Total Pressure at the Outlet of the High Pressure Compressor (HPC) • T30 = HPC Outlet Temperature • P40 = Total Pressure at the Combustion Chamber Outlet • T40 = Temperature at the Combustion Chamber Outlet • P41 = Total Pressure at the Inlet of the High Pressure Turbine (HPT) Rotor • T41 = Temperature at the HPT Rotor Inlet

[0333] In current engines 10, T40 and T41 are 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, the values ​​of T40 and / or T41 can instead be deduced from other measurements (for example, using thermocouple readings used for temperature measurement at other stations and knowledge of the gas turbine engine architecture and thermal properties).

[0334] The relationship between the pressure or temperature values ​​at station 30 and at station 40 or 41 depends on how the motor 10 is controlled / the parameter that is kept constant.

[0335] For example, for an engine 10 operating at a fixed (gravity-driven) fuel flow rate, T41 would generally increase with the introduction of SAF, or a mixture containing more SAF, due to the generally higher calorific value. This change in T41 (or equivalently T40) is then followed by a corresponding increase in shaft speeds and T30 / P30. After the transient changes in the relationship during the fuel type change, the steady-state relationship between T30 and T41 can return to its initial state.

[0336] If, on the other hand, the engine 10 operates with a fixed shaft speed, the fuel mass flow rate decreases when a fuel with a higher calorific value is used, and the core flow rate increases. After the transient changes in the relationship during the change in fuel mass flow rate, the steady-state relationship between T30 and T41 can again return to its initial state.

[0337] In examples where relative temperatures and / or pressures (relationships between temperatures or pressures) are used, a change in the relationship between temperatures and / or pressures over time around the fuel change can be used to infer or calculate one or more fuel characteristics, instead of, or in addition to, examining a ratio of or difference between selected temperatures or pressures at a specific time before the change and at a specific time after the change. Information can therefore be derived from the transient behavior.

[0338] In some examples, the aircraft 1 may have only one fuel tank 50, and / or may have multiple fuel tanks 50, 53, each containing the same fuel, and / or are fluidically connected, or are in fluidic communication with the gas turbine engine 10, so that only one type of fuel is supplied to the gas turbine engine 10 between refueling events—that is, the fuel characteristics may remain constant throughout a flight. In such examples, the change in temperature(s) and / or pressure(s) may therefore be noted on the basis of data recorded for a previous flight (since the last refueling event) or an earlier stage of the same flight relative to the current data, rather than taking pressure and / or temperature data before and after a change occurring during the same flight.Alternatively, relationship data between temperatures and / or pressures for a reference or standard fuel can be provided, and current data can be compared to these. However, it should be understood that, due to the number of potential variables involved and the possibility that some sensor data may not be accurate (e.g., fuel flow rate), it may be preferable to use data obtained just before and after a given change in the described determination (taking into account all transients) and / or during the fuel change (including transient behavior), so as to minimize uncontrolled variables and / or changes in environmental parameters. The examples described herein may therefore be particularly useful in scenarios with at least two fuel sources.

[0339] In such examples, the aircraft 1 may have a plurality of fuel tanks 50, 53 which may contain fuels of different compositions, and the propulsion system 2 may include an adjustable fuel distribution system, allowing the selection of the tank(s) 50, 53, and therefore which fuel / fuel mixture to use. In such examples, the fuel characteristics may vary during a flight. The temperature(s) and / or pressure(s) may be checked each time a fuel change is made, in order to allow the properties of the current fuel to be determined. Alternatively, the temperature(s) and / or pressure(s) may be checked only when switching to a new tank 50, 53 or a new fuel mixture, for which the fuel characteristics have not been previously determined and stored.In such examples, temperature and / or pressure monitoring may be performed at several different times, with an active fuel management system 214 arranged to change the fuel, or fuel mixture, between these times. Fuel characteristics for the multiple different fuels Fb F2 on board. These factors can therefore be determined. The fuel supplied to the gas turbine engine 10 can be changed during cruise, allowing temperature and / or pressure monitoring under relatively constant conditions, so that the fuel change is effectively the only change. This can allow for a more precise determination of any changes in the relationship(s) between temperatures and / or pressures. Similarly, the fuel supplied to the gas turbine engine 10 can be changed while the aircraft is idling on the ground, for example, before takeoff. Again, this can provide relatively constant conditions, so that the fuel change is effectively the only change.

[0340] The temperature(s) and / or pressure(s) can therefore be monitored in two different time periods—one for each of the two different fuels Fb F2, or during a single time period including the fuel change. The fuel change can be the only change made to the engine control between the two time periods / during the single time period. When two separate time periods are used, the two time periods can also be selected so that the altitude and / or other external parameters are at least substantially the same for both, and can therefore be selected 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.Similarly, when a single time period is used, it can be selected to be short enough so that the altitude and / or other external parameters are at least substantially the same throughout the period.

[0341] When changes are evaluated between two distinct time periods, as described above, it may be desirable for the first and second time periods to be as close together as reasonably possible—a small interval may be left to ensure a complete fuel change in the combustion chamber 16 and to allow for any transient effects. (In other implementations, the transient behavior itself may be used to determine the fuel characteristic(s).) The required interval size (if any) may depend on the fuel flow rate under operating conditions. The gas turbine engine 10 typically reacts almost instantaneously (within one second) to fuel differences once that fuel reaches the combustion chamber 16, and the speed probes used for shaft speed measurements generally have a small time constant.Under conditions of relatively low power, low fuel flow, an interval of about ten seconds from the moment the fuel enters the mast that connects the engine 10 to the cell. The time from the aircraft 1 fuel change can be used. At higher power, where the fuel flow rate can be four times greater or more, an interval of 2 to 3 seconds from the fuel change to the mast inlet may be appropriate. It will be understood that the travel time from a fuel tank to the engine 10 can vary based on the tank's location as well as the fuel flow rate, and can be adjusted accordingly based on knowledge of the specific aircraft 1 – the mast inlet is therefore mentioned here for generalization purposes, although the 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 implementations, with the interval calculated with reference to the fuel flow time between the point of interest and the engine 10.

[0342] In addition, the average of the measurements can be established over a period of time (for example, 5 seconds up to 30 seconds) in each time period, or in the second time period only, and any trend can be examined, to verify that a new steady state has been reached and / or to improve reliability.

[0343] Based on knowledge of fuel characteristics, a specific fuel or fuel mixture can be selected to improve performance at certain stages of flight or under certain external conditions.

[0344] Additional data can be used in conjunction with the determined fuel characteristics to adjust the control of the propulsion system 2 and / or changes to the flight profile. For example, the method may include receiving data on the current conditions around the aircraft 1 (either from a supplier, such as a third-party weather monitoring agency, or from onboard sensors). This received data (e.g., weather data, temperature, humidity, contrail presence, etc.) can be used to make or influence changes to the propulsion system control. Instead of, or in addition to, using "real-time" or near-real-time weather data, forecast weather data for the aircraft's route can also be used to estimate current conditions.As used here, the term "flight profile" refers to the operational characteristics (e.g., height / altitude, power setting, flight path angle, aerodynamic speed, and the like) of an aircraft when flying along a flight path, as well as the flight path / route (route) itself. Changes in route (even of only about 100 m) are therefore included in the term "flight profile" as used here.

[0345] Examples of options for controlling the propulsion system 2 based on knowledge of fuel characteristics include those listed above.

[0346] A propulsion system 2 for an aircraft 1 may therefore include a fuel composition tracking device 210 arranged to record and store fuel characteristic data, and optionally also to receive measurement data relating to temperatures and / or pressures inside the gas turbine engine 10, and to determine one or more fuel characteristics on the basis of this data (the determination possibly involving the calculation of a relationship between temperatures and / or pressures between multiple temperatures or pressures, respectively) and optionally other data, such as measurement data relating to responses to one or more operating changes (non-limiting examples of appropriate operating changes are listed above).

[0347] The fuel composition monitoring device 210 may be provided as a separate fuel composition monitoring unit 210 integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems.

[0348] Data from the fuel composition monitoring device 210 can be used to adjust the control of the propulsion system 2, based on the fuel characteristic or characteristics.

[0349] A plurality of temperature and / or pressure sensors 204 may be provided at selected locations in the gas turbine engine 10. In the examples described, multiple sensors are provided for each location of interest, possibly arranged symmetrically around the turbine rotor inlet, for example, so as to provide improved accuracy of the temperature and / or pressure measurements obtained.

[0350] In the example shown, two sensors 204 are provided, each arranged to detect one or more pressures or temperatures related to the performance of the gas turbine engine. The sensors can directly measure one or more of P30, T30, P40, T40, P41, and T41, or can provide other measurements from which one or more of these values ​​can be calculated or deduced. In different implementations, different numbers and / or types of sensors can be provided, as described above.

[0351] The sensors 204 and the fuel composition monitoring device 202 can be described together as a fuel composition monitoring system 203, as shown in [Fig.8], and a fuel composition monitoring system 203 and an EEC 42 can be as described above.

[0352] A method 2010 for determining one or more fuel characteristics of a fuel supplied to a gas turbine engine 10 of an aircraft 1 can therefore be implemented, the gas turbine engine 10 being part of a propulsion system 2.

[0353] The method 2010 includes changing the fuel supplied to a gas turbine engine 10 of an aircraft 1. The change can be carried out during the operation of the aircraft 1 – for example, by using a fuel management system 214 to draw fuel from a different tank 50, 53 – or between different operating sessions of an aircraft 1 – for example, when refueling an aircraft 1 with new fuel. The fuel change can be temporary and can be reversed as soon as sufficient time has elapsed for any effect on the temperature(s) and / or pressure(s) to be detected.

[0354] The method 2010 further includes the detection 2014 of a response to a change in fuel, and in particular the detection, determination, or inference of a change in at least one selected temperature and / or pressure(s). Optionally, two or more temperatures or pressures may be detected, so that a relationship between P30 and one or more of P41 and P40, or T30, and one or more of T41 and T40, may be determined on the basis of the sensor data. For example, a change in one or more of the listed pressures and / or temperatures may be detected directly or inferred / determined / calculated from other measurements and knowledge of the engine 10.

[0355] The 2010 method further includes determining 2016 one or more fuel characteristics of the fuel burned by the gas turbine engine 10 based on the response to the fuel change. For example, a percentage change in calorific value between the first fuel (before the change) and the second fuel can be determined so as to provide knowledge of the relative fuel properties, and / or an actual calorific value can be determined (either directly or by using knowledge of the values ​​for the first fuel).

[0356] The 2012 fuel change, and subsequent steps of the 2010 process, can be repeated to confirm the fuel characteristics obtained.

[0357] In some implementations, the 2010 procedure may further include making one or more changes to the aircraft operation and / or a planned flight profile after the 2016 determination is made, based on the determined fuel characteristic(s), for example, to improve engine efficiency or reduce the impact on the climate (e.g., by adjusting the formation of condensation trails). In other implementations, knowledge of the fuel characteristics may not be used to change the aircraft operation, but may be used to influence refueling choices and / or to verify that the fuel data provided for a fuel is correct. In the event of a significant discrepancy between the determined fuel characteristics and the expected fuel characteristics, the aircraft may The aircraft may be returned to a refueling station for fuel verification, and / or further checks may be carried out. EEC 42 can be configured to provide a warning / alert to a pilot in such scenarios.

[0358] In implementations in which a fuel composition monitoring device 202, 210 as described above is used to carry out part or all of process 2010, the fuel composition monitoring device 202, 210 can be arranged to receive data corresponding to a change in one or more of T30, P30, T40, T41, P40 and P41; and determine one or more fuel characteristics of the fuel based on the change in temperature(s) and / or pressure(s).

[0359] In certain cases, the fuel composition monitoring device 202, 210 can be arranged to:

[0360] receive data corresponding to a change in a relationship between T30 (or P30) and one of T40 and T41 (or one of P40 and P41); and

[0361] determine one or more fuel characteristics of the fuel based on the change in the relationship between temperatures and / or pressures.

[0362] In examples with two or more fuel sources, the propulsion system 2 may further include a fuel management system, for example a fuel management device 214, arranged to change the fuel supplied to the gas turbine engine 10 in flight; for example by actively selecting a particular tank 50, 53, or a particular fuel mixture from among several tanks, in flight. A propulsion system control device (for example the EEC 42) may be used to adjust the control of the propulsion system 2 based on the fuel characteristic(s) of the fuel, based on data provided by the fuel composition tracking device 202 and possibly other data.The propulsion system control device 42 may be provided as a separate propulsion system control unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems. The fuel composition monitoring means may be provided as part of the same unit or assembly.

[0363] As described above, the propulsion system control device 42 can make changes to the propulsion system directly, or can provide a notification to the pilot recommending the change for approval. In some examples, the same propulsion system control device 42 can automatically perform some changes and request others, depending on the nature of the change, as discussed above.

[0364] The propulsion system control device 42 can also provide Recommendations concerning flight profile changes. Alternatively, or in addition, Propulsion System 2 may therefore include a flight profile adjustment device arranged to change the planned flight profile based on the fuel's fuel characteristics, and possibly other data. The flight profile adjustment device may be provided as a separate propulsion system control unit integrated into Propulsion System 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems such as EEC 42. Fuel composition monitoring means may be provided as part of the same unit or assembly.

[0365] It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements may be made without departing from the concepts described herein. Except as mutually exclusive, any of the features may be used separately or in combination with any other feature, and disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Demands

1. A method (2090) for determining at least one fuel characteristic of a fuel supplied to a gas turbine engine (10) of an aircraft (1), the gas turbine engine (10) being part of a propulsion system (2), the method comprising: effecting (2092) a change in operation, the change in operation being effected by a controllable component of the propulsion system (2) and arranged to affect the operation of the gas turbine engine (10); detecting (2094) a response to the change in operation; and determining (2096) at least one fuel characteristic on the basis of the response to the change in operation.

2. Method (2090) of claim 1, wherein the propulsion system (2) includes a thermal management system, and wherein the implementation (2092) of a change of operation includes the use of the thermal management system to change the temperature of the fuel entering a combustion chamber of the gas turbine engine (10).

3. Method (2090) of claim 2, wherein the response to the change in operation includes at least one of: (i) a change in the power output of the gas turbine engine (10); (ii) a change in the degradation or coking of the fuel; (iii) a change in at least one pressure in the engine (10); and (iv) a change in at least one temperature in the engine (10).

4. Method (2090) of claim 1, wherein the propulsion system (2) comprises at least one steerable inlet guide vane -VIGV, and wherein the implementation (2092) of a change of operation comprises the programming change of VIGV.

5. Method (2090) of claim 4, wherein the response to the change in operation includes at least one of: (i) a change in the gas temperature at the inlet of a turbine of the gas turbine engine (10); and (ii) a temperature rise through a combustion chamber (16) of the gas turbine engine (10).

6. A method (2090) of claim 1, wherein the system of propulsion (2) comprises a plurality of fuel tanks (50, 53), and wherein the realization (2092) of a change of operation includes the change of the tank (50, 53) from which the fuel is taken.

7. A method (2090) of claim 1, wherein the propulsion system (2) comprises a plurality of fuel tanks (50, 53), and wherein the implementation (2092) of a change of operation comprises changing the percentage of fuel taken from a particular tank.

8. Method (2090) of claim 6 or 7, wherein the response to the change in operation includes at least one of: (i) a change in the power output of the gas turbine engine (10); (ii) a change in the degradation or coking of the fuel; (iii) a change in the formation of condensation trail; (iv) a change in the relationship between a temperature at the compressor outlet and a temperature at the turbine rotor inlet; (v) a change in the relationship between a total pressure at the compressor outlet and a total pressure at the turbine rotor inlet.

9. Method (2090) of claim 1, wherein the propulsion system (2) comprises at least one air-oil heat exchanger (118), and wherein the implementation (2092) of a change of operation comprises changing at least one of the air flow and oil flow through the air-oil heat exchanger (118).

10. Method (2090) of claim 9, wherein the response to the change in operation includes a change in pressure in a fuel circuit (3) of the gas turbine engine (10).

11. A method (2090) of claim 1, wherein at least one fuel characteristic comprises at least one of the following: i. the percentage of sustainable aviation fuel in the fuel; ii. the aromatic hydrocarbon content of the fuel; iii. the multi-aromatic hydrocarbon content of the fuel; iv. the percentage of nitrogen-containing species in the fuel; v. the presence or percentage of a tracer or trace substance in the fuel; vi. the hydrogen-to-carbon ratio of the fuel; vii. the hydrocarbon distribution of the fuel; viii. the level of non-volatile particulate matter emissions during combustion; ix. the naphthalene content of the fuel; x. the sulfur content of the fuel; xi. the cycloparaffin content of the fuel; xii. the oxygen content of the fuel; xiii. the thermal stability of the fuel; xiv. the level of coking of the fuel; xv. an indication that the fuel is a fossil fuel; and xvi. at least one of the following: density, viscosity, calorific value, and heat capacity.

12. Propulsion system (2) for an aircraft (1) comprising: a gas turbine engine (10); a fuel tank (50, 52) arranged to contain fuel to supply the gas turbine engine (10); and a fuel composition tracking device (202) arranged to: receive information concerning a change in operation, the change in operation being effected by a controllable component of the propulsion system (2) and arranged to affect the operation of the gas turbine engine (10); receive data corresponding to a response to the change in operation; and determine at least one fuel characteristic of the fuel arranged to be supplied to the gas turbine engine (10) based on the response to the change in operation.

13. Propulsion system (2) of claim 12, further comprising at least one sensor (204) arranged to detect the response to the change in operation and to provide data to the fuel composition monitoring device (202).

14. Propulsion system (2) of claim 12, wherein at least one sensor (204) comprises at least one of: (i) a temperature sensor; and (ii) a pressure sensor.

15. Propulsion system (2) of claim 12, further comprising at least one steerable inlet guide vane - VIGV, and wherein the change of operation includes the change of programming of VIGV.

16. Propulsion system (2) of claim 12, comprising a plurality of fuel tanks (50, 53), and wherein the change of operation includes changing the tank (50, 53) from which the fuel is taken.

17. Propulsion system (2) of claim 12, comprising a plurality of fuel tanks (50, 53), and wherein the change of operation includes changing the percentage of fuel taken from a particular tank.

18. Propulsion system (2) of claim 12, further comprising at least one air-oil heat exchanger (118), and wherein the change of operation comprises the change of at least one of the air flow and the oil flow through the air-oil heat exchanger.

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