Fuel flow moderation in a gas turbine engine

Abstract

A gas turbine engine (10) for an aircraft (1) comprises: a combustion chamber (16); a fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) arranged to receive fuel from a fuel tank (50) on board the aircraft (1) and transfer heat from the oil to the fuel; a fuel return line (1011; 1211; 1411b; 1711b) for returning at least a portion of the fuel that has passed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulating valve (1010; 1210; 1410; 1710) arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b;1711b) such that a ratio between ΔT at cruise and ΔT at startup is less than 0.56, wherein ΔT is a ratio between a temperature, in kelvin, of the fuel in the fuel tank and a temperature, in kelvin, of the fuel delivered to the combustion chamber. The invention also discloses a method (2100) of operating a gas turbine engine. Figure for abstract: [Figure 5];

Description

Title of the invention: Moderation of fuel flow in a gas turbine engine

[0001] The present disclosure relates to gas turbine engines designed to operate using fuels other than traditional kerosene-based jet fuels, and methods of operating a gas turbine engine using fuels other than traditional kerosene-based jet fuels.

[0002] There is an expectation in the aviation industry regarding a trend towards the use of fuels other than the traditional kerosene-based jet fuels generally used today. It is desirable to manage the heat and fuel within the engine according to the type of fuel used. This may allow advantageous use of fuel properties that are not available for kerosene-based jet fuel.

[0003] According to a first aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0004] a combustion chamber;

[0005] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0006] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0007] a modulating valve for modulating the flow of fuel along the fuel return line such that a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.56.

[0008] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.53.

[0009] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.56.

[0010] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.53.

[0011] According to a second aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0012] a combustion chamber;

[0013] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0014] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0015] a modulating valve for modulating the flow of fuel along the fuel return line;

[0016] wherein the method comprises modulating the flow of fuel along the fuel return line using the modulating valve such that a ratio of a temperature, in kelvin, of the fuel in the tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.56.

[0017] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.53.

[0018] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.56.

[0019] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.53.

[0020] According to a third aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0021] a combustion chamber;

[0022] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0023] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0024] a modulating valve for modulating the flow of fuel along the fuel return line such that a ratio between a temperature, in kelvin, of the fuel in the fuel tank and a temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 1.00.

[0025] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is:

[0026] a) between 0.48 and 0.90;

[0027] b) between 0.48 and 0.82;

[0028] c) between 0.56 and 1.00;

[0029] d) between 0.56 and 0.82; or

[0030] e) between 0.56 and 0.75.

[0031] According to a fourth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0032] a combustion chamber;

[0033] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0034] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0035] a modulating valve for modulating the flow of fuel along the fuel return line;

[0036] wherein the method comprises modulating the flow of fuel along the fuel return line using the modulating valve such that a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 1.00.

[0037] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is:

[0038] a) between 0.48 and 0.90;

[0039] b) between 0.48 and 0.82;

[0040] c) between 0.56 and 1.00;

[0041] d) between 0.56 and 0.82; or

[0042] e) between 0.56 and 0.75.

[0043] According to a fifth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0044] a combustion chamber;

[0045] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0046] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0047] a modulating valve for modulating fuel flow along the fuel return line such that a ratio of AT at cruise to AT at startup is less than 0.56, wherein AT is a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber.

[0048] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of AT at cruise to AT at startup is less than 0.53.

[0049] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of AT at cruise to AT at startup is between 0.48 and 0.56.

[0050] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of AT at cruise to AT at startup is between 0.48 and 0.53.

[0051] According to a sixth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0052] a combustion chamber;

[0053] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0054] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0055] a modulating valve for modulating fuel flow along the fuel return line, wherein the method comprises modulating fuel flow along the fuel return line using the modulating valve such that a ratio of AT at cruise to AT at startup is less than 0.56, wherein AT is a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber.

[0056] The method may include modulating the fuel flow along the fuel return line using the modulating valve such that the ratio of AT at cruise to AT at startup is less than 0.53.

[0057] The method may include modulating the fuel flow along the fuel return line using the modulating valve such that the ratio of AT at cruise to AT at startup is between 0.48 and 0.56.

[0058] The method may include modulating the fuel flow along the fuel return line using the modulating valve such that the ratio of AT at cruise to AT at startup is between 0.48 and 0.53.

[0059] According to a seventh aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0060] a combustion chamber;

[0061] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0062] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0063] a modulating valve for modulating fuel flow along the fuel return line such that a ratio of AT at cruise to AT at startup is between 0.48 and 1.88, wherein AT is a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber.

[0064] The modulating valve may be intended to modulate the flow of fuel along the fuel return line such that the ratio between AT at cruise and AT at startup is:

[0065] a) between 0.48 and 1.71;

[0066] b) between 0.48 and 1.70 (for example, where AT at startup is AT at startup of the aircraft when it is on the ground);

[0067] c) between 0.48 and 1.55 (for example, where AT at start is AT at start of the aircraft when it is on the ground);

[0068] d) between 0.48 and 1.88 (e.g., where AT at startup is AT at startup of the aircraft when in flight); or

[0069] e) between 0.48 and 1.71 (where AT at start is AT at start of the aircraft when it is in flight).

[0070] According to an eighth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0071] a combustion chamber;

[0072] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0073] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0074] a modulating valve for modulating fuel flow along the fuel return line, wherein the method comprises modulating fuel flow along the fuel return line using the modulating valve such that a ratio between AT at cruise and AT at startup is between 0.48 and 1.88, where AT is a ratio between a temperature, in kelvin, of the fuel in the fuel tank and a temperature, in kelvin, of the fuel delivered to the combustion chamber.

[0075] The method may include modulating the fuel flow along the fuel return line using the modulating valve such that the ratio of AT at cruise to AT at startup is:

[0076] a) between 0.48 and 1.71;

[0077] b) between 0.48 and 1.70 (for example, where AT at startup is AT at startup of the aircraft when it is on the ground);

[0078] c) between 0.48 and 1.55 (for example, where AT at start is AT at start of the aircraft when it is on the ground);

[0079] d) between 0.48 and 1.88 (e.g., where AT at startup is AT at startup of the aircraft when in flight); or

[0080] e) between 0.48 and 1.71 (where AT at start is AT at start of the aircraft when it is in flight).

[0081] According to a ninth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0082] a combustion chamber;

[0083] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0084] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0085] a modulating valve for modulating the flow of fuel along the fuel return line, wherein the modulating valve is for initiating a return of fuel to the fuel tank when the fuel having passed through the heat exchanger is at a temperature of at least 120°C.

[0086] The modulating valve may be intended to initiate a return of fuel to the fuel tank when the fuel having passed through the heat exchanger is at a temperature of at least 140°C.

[0087] The modulating valve may be intended to initiate a return of fuel to the fuel tank when the fuel having passed through the heat exchanger is at a temperature between 120°C and 180°C, and preferably at a temperature between 140°C and 180°C.

[0088] The modulating valve may be intended to initiate a return of fuel to the fuel tank when: i. an indication of an operating condition is provided; and ii) the fuel having passed through the heat exchanger is at a temperature of at least 120°C, preferably at least 140°C.

[0089] The operating conditions may be one or more of a proportion of sustainable aviation fuel (SAF) in the fuel, a thermal stability of the fuel, a coking level of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0090] The gas turbine engine may further include a sensor configured to detect one or more operating conditions.

[0091] According to a tenth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0092] a combustion chamber;

[0093] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0094] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0095] a modulating valve for modulating the flow of fuel along the fuel return line;

[0096] wherein the method comprises initiating a return of fuel to the fuel tank using the modulating valve when the fuel having passed through the heat exchanger is at a temperature of at least 120°C.

[0097] The method may include initiating a return of fuel to the fuel tank using the modulating valve when the fuel having passed through the heat exchanger is at a temperature of at least 140°C.

[0098] The method may comprise initiating a return of fuel to the fuel tank using the modulating valve when the fuel having passed through the heat exchanger is at a temperature between 120°C and 180°C, and preferably at a temperature between 140°C and 180°C.

[0099] The method may include initiating a return of fuel to the fuel tank using the modulating valve when: i. an indication of an operating condition is provided; and ii) the fuel having passed through the heat exchanger is at a temperature of at least 120°C, preferably at least 140°C.

[0100] The operating conditions may be one or more of a proportion of sustainable aviation fuel (SAF) in the fuel, a thermal stability of the fuel, a coking level of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0101] The gas turbine engine may further include a sensor configured to detect one or more operating conditions. The method may include detecting one or more operating conditions using the sensor.

[0102] According to an eleventh aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0103] a combustion chamber;

[0104] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0105] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0106] a modulating valve for modulating the flow of fuel along the fuel return line such that a ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruising conditions.

[0107] The "mass of fuel returned to the fuel tank" and the "mass of fuel delivered to the combustion chamber" may be expressed as a fuel flow rate in fuel flow mass per unit time (e.g., kg / s). By "mass of fuel delivered to the combustion chamber" in the eleventh and twelfth aspects, it may be meant the "mass of fuel burned", e.g., the amount of fuel burned by the combustion chamber in units of fuel flow rate per unit time.

[0108] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0109] The modulating valve may be intended to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0110] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0111] According to a twelfth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0112] a combustion chamber;

[0113] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0114] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0115] a modulating valve for modulating the flow of fuel along the fuel return line;

[0116] wherein the method comprises modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the mass of fuel returned to the fuel tank to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruising conditions.

[0117] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that a ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0118] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that a ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0119] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0120] According to a thirteenth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0121] a combustion chamber;

[0122] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0123] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0124] a modulating valve for modulating the flow of fuel along the fuel return line, wherein the modulating valve is for preventing fuel having a temperature of 180°C or higher from being returned to the fuel tank.

[0125] The modulating valve may be adapted to prevent fuel from returning to the fuel tank when a temperature of the fuel in the fuel tank is at a predetermined upper threshold temperature.

[0126] The predetermined upper threshold temperature may be 100°C. The predetermined upper threshold temperature may be 65°C. The upper threshold temperature predetermined can be 55°C. The predetermined upper threshold temperature can be 5°C.

[0127] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0128] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that an equilibrium temperature of the fuel in the fuel tank, when mixed with the fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature.

[0129] According to a fourteenth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0130] a combustion chamber;

[0131] a fuel-oil heat exchanger for receiving fuel from a fuel tank on board the aircraft and transferring heat from the oil to the fuel;

[0132] a fuel return line for returning at least a portion of the fuel that has passed through the heat exchanger to the fuel tank; and

[0133] a modulating valve for modulating the flow of fuel along the fuel return line;

[0134] wherein the method comprises preventing return of fuel having a temperature of 180°C or more to the fuel tank using the modulating valve.

[0135] The method may include preventing a return of fuel to the fuel tank using the modulating valve when a temperature of the fuel in the fuel tank is at a predetermined upper threshold temperature.

[0136] The predetermined upper threshold temperature may be 100°C. The predetermined upper threshold temperature may be 65°C. The predetermined upper threshold may be 55°C. The predetermined upper threshold temperature may be 5°C.

[0137] The method may include modulating a flow of fuel along the fuel return line using the modulating valve based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0138] The method may include modulating a flow of fuel along the fuel return line using the modulating valve such that a equilibrium temperature of the fuel in the fuel tank, when mixed with the fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature.

[0139] According to a fifteenth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising:

[0140] a combustion chamber;

[0141] a fuel-oil heat exchanger for receiving the fuel and transferring heat from the oil to the fuel so as to raise the temperature of the fuel to at least 120°C at the inlet of the combustion chamber;

[0142] a fuel recirculation line for recirculating at least a portion of fuel along a fuel flow path from a first point in the fuel flow path to a second point in the fuel flow path, the second point being upstream of the first point; and

[0143] a modulating valve for modulating the flow of fuel along the fuel recirculation line.

[0144] The fuel recirculation line may be adapted to recirculate fuel from a first point downstream of the fuel-oil heat exchanger to a second point upstream of the fuel-oil heat exchanger.

[0145] The gas turbine engine may also include a fuel pump in the fuel flow path. The fuel recirculation line may be to recirculate fuel from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

[0146] The fuel pump may be located downstream of the fuel-oil heat exchanger in the fuel flow path.

[0147] The fuel recirculation line may be for supplying fuel to one or more additional aircraft and / or engine mechanisms.

[0148] The additional aircraft and / or engine mechanism(s) may be or include one or more of a nacelle anti-icing system, actuators, purge valves, heat management modulating valves (e.g., for the engine and / or generator system and a turbine case cooling system).

[0149] The fuel flow path may be a primary fuel flow path between a fuel tank on board the aircraft and the combustion chamber. The fuel-oil heat exchanger may be a primary fuel-oil heat exchanger located on the primary fuel flow path.

[0150] The second point may be located downstream of a low pressure fuel pump by which fuel is pumped to the gas turbine engine from the fuel tank.

[0151] The first point may be located downstream of the primary fuel-oil heat exchanger. The first point may be located downstream of a fuel pump configured to supply fuel to the combustion chamber (i.e., a high-pressure fuel pump).

[0152] The engine may further include a secondary fuel-oil heat exchanger located in the primary fuel path (e.g., in series with the primary heat exchanger). The second point may be located downstream of the secondary heat exchanger.

[0153] Alternatively, the engine may further comprise a secondary fuel-oil heat exchanger with a flow line leading from the primary flow path to the secondary fuel-oil heat exchanger and a flow line joining the primary flow path from the secondary heat exchanger (e.g., such that it is parallel to the primary flow path). The second point may be located upstream of the inlet of the line leading to the secondary fuel-oil heat exchanger and the first point may be located downstream of an outlet of the line joining the primary fuel flow path from the secondary fuel-oil heat exchanger. The second point may be located upstream of the primary fuel-oil heat exchanger.

[0154] The fuel-to-oil heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to at least 140°C at the inlet to the combustion chamber (i.e., the heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to these temperatures before it enters the combustion chamber).

[0155] The heat exchanger may be intended to transfer heat from the oil to the fuel in order to raise the temperature of the fuel to between 120°C and 180°C, and preferably between 140°C and 180°C, at the inlet of the combustion chamber.

[0156] The heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to at least 120°C, 125°C, 130°C, 135°C, 140°C, 145°C, 150°C, 155°C, 160°C, 165°C, 170°C, 175°C, 180°C, 185°C, 190°C, 195°C or 200°C at the inlet to the combustion chamber, or to a fuel temperature in a defined range between any two of these values.

[0157] The modulating valve may be adapted to modulate the flow of fuel along the fuel recirculation line such that a ratio of the mass of recirculated fuel to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruising conditions.

[0158] The modulating valve may be adapted to modulate the flow of fuel along the fuel recirculation line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0159] The modulating valve may be intended to modulate the flow of fuel along the fuel recirculation line such that the ratio between the mass of fuel returned to the tank and the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0160] The modulating valve may be adapted to modulate the flow of fuel along the fuel recirculation line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0161] The gas turbine engine may further comprise at least one temperature sensor on the fuel flow path and / or on the fuel recirculation line.

[0162] The modulating valve may be adapted to modulate the flow of fuel along the fuel recirculation line based, at least in part, on temperature data from the at least one temperature sensor.

[0163] According to a sixteenth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0164] a combustion chamber;

[0165] a fuel-oil heat exchanger for receiving fuel from a fuel tank and transferring heat from the oil to the fuel so as to raise the temperature of the fuel to at least 120°C at the inlet of the combustion chamber;

[0166] a fuel recirculation line for recirculating at least a portion of fuel along a fuel flow path from a first point in the fuel flow path to a second point in the fuel flow path, the second point being upstream of the first point; and

[0167] a modulating valve for modulating the flow of fuel along the fuel recirculation line;

[0168] wherein the method comprises modulating the flow of fuel along the fuel recirculation line using the modulating valve.

[0169] The method may include modulating the flow of fuel along the fuel recirculation line, using the modulating valve, from a first point downstream of the fuel-oil heat exchanger to a second point upstream of the fuel-oil heat exchanger.

[0170] The gas turbine engine may also include a fuel pump in the fuel flow path. The method may include modulating the flow of fuel along the fuel recirculation line, using the modulating valve, from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

[0171] The fuel pump may be located downstream of the fuel-oil heat exchanger in the fuel flow path.

[0172] The method may include supplying fuel to one or more additional aircraft and / or engine mechanisms via the fuel recirculation line.

[0173] The additional aircraft and / or engine mechanism(s) may be or include one or more of a nacelle anti-icing system, actuators, purge valves, heat management modulating valves (e.g., for the engine and / or generator system and a turbine case cooling system.

[0174] The fuel flow path may be a primary flow path between the fuel tank on board the aircraft and the combustion chamber. The fuel-oil heat exchanger may be a primary fuel-oil heat exchanger located on the primary fuel flow path.

[0175] The second point may be located downstream of a low pressure fuel pump by which fuel is pumped to the gas turbine engine from the fuel tank.

[0176] The first point may be located downstream of the primary fuel-oil heat exchanger. The first point may be located downstream of a fuel pump configured to supply fuel to the combustion chamber (i.e., a high-pressure fuel pump).

[0177] The engine may further include a secondary fuel-oil heat exchanger located in the primary fuel path (e.g., in series with the primary heat exchanger). The second point may be located downstream of the secondary heat exchanger.

[0178] Alternatively, the engine may further comprise a secondary fuel-oil heat exchanger with a flow line leading from the primary flow path to the secondary fuel-oil heat exchanger and a flow line joining the primary flow path from the secondary heat exchanger (e.g., such that it is parallel to the primary flow path). The second point may be located upstream of the inlet of the line leading to the secondary fuel-oil heat exchanger and the first point may be located downstream of an outlet of the line joining the primary fuel flow path from the fuel-oil heat exchanger secondary. The second point can be located upstream of the primary fuel-oil heat exchanger.

[0179] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature to at least 140°C at the inlet to the combustion chamber.

[0180] The method comprises transferring heat from the oil to the fuel so as to raise the fuel temperature to between 120°C and 180°C at the inlet into the combustion chamber.

[0181] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature to between 140°C and 180°C at the inlet into the combustion chamber.

[0182] The heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to at least 120°C, 125°C, 130°C, 135°C, 140°C, 145°C, 150°C, 155°C, 160°C, 165°C, 170°C, 175°C, 180°C, 185°C, 190°C, 195°C or 200°C at the inlet to the combustion chamber, or to a fuel temperature in a defined range between any two of these values.

[0183] The method may include modulating the flow of fuel along the recirculation line using the modulating valve such that a ratio of the mass of fuel recirculated to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0184] The method may include modulating the flow of fuel along the recirculation line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0185] The method may include modulating the flow of fuel along the recirculation line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0186] The method may include modulating the flow of fuel along the recirculation line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0187] The gas turbine engine may further comprise at least one temperature sensor on the fuel flow path and / or on the fuel recirculation line.

[0188] The method may include modulating the flow of fuel along the fuel recirculation line, using the modulating valve, based at least in part on temperature data from the at least one temperature sensor.

[0189] The fuel recirculation features of the fifteenth and sixteenth aspects may be combined with any of the other aspects defined above or elsewhere herein. In other words, fuel may be recirculated back to the fuel tank using the modulating valve according to any aspect defined above or elsewhere herein, as well as recirculated as defined in the fifteenth and sixteenth aspects.

[0190] In the fifth, seventh, ninth, eleventh and thirteenth aspects:

[0191] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that a ratio of the temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.56 (e.g., at cruise conditions).

[0192] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.53 (e.g., at cruise conditions).

[0193] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.56 (e.g., at cruise conditions).

[0194] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.53 (e.g., at cruise conditions).

[0195] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 1.00 (e.g., at cruise conditions).

[0196] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber (e.g., under cruising conditions) is:

[0197] a) between 0.48 and 0.90;

[0198] b) between 0.48 and 0.82;

[0199] c) between 0.56 and 1.00;

[0200] d) between 0.56 and 0.82; or

[0201] e) between 0.56 and 0.75.

[0202] In the sixth, eighth, tenth, twelfth and fourteenth aspects:

[0203] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that a ratio of a temperature, in kelvin, of the fuel in the fuel tank to a temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.56 (e.g., at cruise conditions).

[0204] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is less than 0.53 (e.g., at cruise conditions).

[0205] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.56 (e.g., at cruise conditions).

[0206] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 0.53 (e.g., at cruise conditions).

[0207] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvin, of the fuel in the fuel tank to the temperature, in kelvin, of the fuel delivered to the combustion chamber is between 0.48 and 1.00 (e.g., at cruise conditions).

[0208] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the temperature, in kelvins, of the fuel in the fuel tank to the temperature, in kelvins, of the fuel delivered to the combustion chamber (e.g., at cruise conditions) is:

[0209] a) between 0.48 and 0.90;

[0210] b) between 0.48 and 0.82;

[0211] c) between 0.56 and 1.00;

[0212] d) between 0.56 and 0.82; or

[0213] e) between 0.56 and 0.75.

[0214] In the first, third, fifth, seventh, eleventh and thirteenth aspects:

[0215] The modulating valve may be intended to initiate a return of fuel to the fuel tank when the fuel having passed through the heat exchanger is at a temperature of at least 120°C and preferably at least 140°C.

[0216] The modulating valve may be intended to initiate a return of fuel to the fuel tank when the fuel having passed through the heat exchanger is at a temperature between 120°C and 180°C, and preferably is at a temperature between 140°C and 180°C.

[0217] The modulating valve may be intended to initiate a return of fuel to the fuel tank when: i. an indication of an operating condition is provided; and ii) the fuel having passed through the heat exchanger is at a temperature of at least 120°C, preferably at least 140°C.

[0218] The operating conditions may be one or more of a proportion of sustainable aviation fuel (SAF) in the fuel, a thermal stability of the fuel, a coking level of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0219] The gas turbine engine may further include a sensor configured to detect one or more operating conditions.

[0220] In the second, fourth, sixth, eighth, twelfth and fourteenth aspects:

[0221] The method may comprise initiating a return of fuel to the fuel tank using the modulating valve when the fuel having passed through the heat exchanger is at a temperature of at least 120°C and preferably at least 140°C.

[0222] The method may comprise initiating a return of fuel to the fuel tank using the modulating valve when the fuel having passed through the heat exchanger is at a temperature between 120°C and 180°C, preferably is at a temperature between 140°C and 180°C.

[0223] The method may include initiating a return of fuel to the fuel tank using the modulating valve when:

[0224] i) an indication of an operating condition is provided; and

[0225] ii) the fuel having passed through the heat exchanger is at a temperature of at least 120°C, preferably at least 140°C.

[0226] The operating conditions may be one or more of a proportion of sustainable aviation fuel (SAF) in the fuel, a thermal stability of the fuel, a coking level of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0227] The gas turbine engine may further include a sensor configured to detect one or more operating conditions. The method may include detecting one or more operating conditions using the sensor.

[0228] In the first, third, fifth, seventh, ninth and thirteenth aspects

[0229] The modulating valve may be intended to modulate the flow of fuel along of the fuel return line such that the ratio of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruising conditions.

[0230] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0231] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0232] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0233] In the second, fourth, sixth, eighth, tenth, fourteenth aspects:

[0234] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the mass of fuel returned to the fuel tank to the mass of fuel delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0235] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is between 2.3 and 9 under cruising conditions.

[0236] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio between the mass of fuel returned to the tank and the mass of fuel distributed to the combustion chamber is greater than 4 and less than or equal to 9 under cruising conditions.

[0237] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruising conditions.

[0238] In the first, third, fifth, seventh, ninth and eleventh aspects:

[0239] The modulating valve may be intended to prevent a return of fuel having a temperature of 180°C or more to the fuel tank.

[0240] The modulating valve may be adapted to prevent fuel from returning to the fuel tank when a temperature of the fuel in the fuel tank is at a predetermined upper threshold temperature.

[0241] The predetermined upper threshold temperature may be 100°C. The predetermined upper threshold temperature may be 65°C. The predetermined upper threshold temperature may be 55°C. The predetermined upper threshold temperature may be 5°C.

[0242] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0243] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line such that an equilibrium temperature of the fuel in the fuel tank, when mixed with the fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature.

[0244] In the second, fourth, sixth, eighth, tenth and twelfth aspects:

[0245] The method may include preventing return of fuel having a temperature of 180°C or higher to the fuel tank using the modulating valve.

[0246] The method may include preventing a return of fuel to the fuel tank using the modulating valve when a temperature of the fuel in the fuel tank is at a predetermined upper threshold temperature.

[0247] The predetermined upper threshold temperature may be 100°C. The predetermined upper threshold temperature may be 65°C. The predetermined upper threshold may be 55°C. The predetermined upper threshold temperature may be 5°C.

[0248] The method may include modulating a flow of fuel along the fuel return line using the modulating valve based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0249] The method may include modulating a flow of fuel along the fuel return line using the modulating valve such that an equilibrium temperature of the fuel in the fuel tank, when mixed with the fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature.

[0250] In the first, second, third and fourth aspects:

[0251] A temperature of the fuel delivered to the combustion chamber (i.e., at the inlet of the combustion chamber) may be a temperature of the fuel delivered to the combustion chamber under cruise conditions. The temperatures of fuel delivered to the combustion chamber under cruise conditions may be defined as an average over at least 5 minutes, 10 minutes, or 30 minutes, under steady-state cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in operating fuel temperature, often a rise in temperature. Each fluctuation may not last more than 5 minutes.

[0252] The preceding paragraph may also apply to other references to the fuel temperature at the combustion chamber inlet.

[0253] In the first, third, fifth, seventh, ninth, eleventh, thirteenth and fifteenth aspects:

[0254] The heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the fuel temperature to at least 120°C, and preferably to a temperature of 140°C, at the inlet to the combustion chamber (i.e., the heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel so that the fuel has these temperatures when it enters the combustion chamber).

[0255] The heat exchanger may be intended to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to between 120°C and 180°C, and preferably between 140°C and 180°C, at the inlet of the combustion chamber.

[0256] The heat exchanger may be adapted to transfer heat from the oil to the fuel so as to raise the temperature of the fuel to at least 125°C, 130 °C, 135 °C, 140 °C, 145 °C, 150 °C, 155 °C, 160 °C, 165 °C, 170 °C, 175 °C, 180 °C, 185 °C, 190 °C, 195 °C or 200 °C at the combustion chamber inlet, or at a fuel temperature within a defined range between any two of these values.

[0257] In the second, fourth, sixth, eighth, tenth and twelfth, fourteenth and sixteenth aspects:

[0258] The method may include transferring heat from the oil to the fuel using the fuel-oil heat exchanger so as to raise the fuel temperature to at least 120°C, and preferably to at least 140°C at the inlet to the combustion chamber.

[0259] The method may comprise transferring heat from the oil to the fuel using the fuel-oil heat exchanger so as to raise the temperature of the fuel to a value between 120°C and 180°C, and preferably between 140°C and 180°C, at the inlet of the combustion chamber.

[0260] The method may comprise transferring heat from the oil to the fuel so as to raise the fuel temperature to at least 125°C, 130°C, 135°C, 140°C, 145°C, 150°C, 155°C, 160°C, 165°C, 170°C, 175°C, 180°C, 185°C, 190°C, 195°C or 200°C at the inlet to the combustion chamber, or to a fuel temperature in a defined range between any two of these values.

[0261] In the first, third, fifth, seventh aspects:

[0262] The gas turbine engine may include a first temperature sensor located downstream of the heat exchanger; and may be configured to receive information from a second temperature sensor located in the fuel tank.

[0263] The first temperature sensor may be located in the fuel return line.

[0264] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based on temperature data from the first temperature sensor and the second temperature sensor.

[0265] In the ninth aspect:

[0266] The gas turbine engine may further comprise at least one temperature sensor located downstream of the heat exchanger.

[0267] The modulating valve may be adapted to initiate a return of fuel to the fuel tank based, at least in part, on temperature data from the at least one temperature sensor located downstream of the heat exchanger.

[0268] In the eleventh and thirteenth aspect s j.

[0269] The gas turbine engine may further comprise at least one temperature sensor located downstream of the heat exchanger; and / or the gas turbine engine may be designed to receive information from a temperature sensor located in the fuel tank.

[0270] The at least one temperature sensor located downstream of the heat exchanger may be located in the fuel return line.

[0271] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based, at least in part, on temperature data from the at least one temperature sensor located downstream of the heat exchanger and / or the temperature sensor located in the fuel tank.

[0272] In the second, fourth, sixth, eighth aspects:

[0273] The gas turbine engine may include a first temperature sensor located downstream of the heat exchanger; and may be configured to receive information from a second temperature sensor located in the fuel tank.

[0274] The first temperature sensor may be located in the fuel return line.

[0275] The method may include modulating the flow of fuel along the fuel return line based on temperature data from the first temperature sensor and the second temperature sensor.

[0276] In the tenth aspect:

[0277] The gas turbine engine may further comprise at least one temperature sensor located downstream of the heat exchanger.

[0278] The method may include initiating a return of fuel to the fuel tank using the modulating valve, based at least in part on temperature data from the at least one temperature sensor located downstream of the heat exchanger.

[0279] In the twelfth and fourteenth aspects:

[0280] The gas turbine engine may further comprise at least one temperature sensor located downstream of the heat exchanger; and / or the gas turbine engine may be configured to receive information from a temperature sensor located in the fuel tank.

[0281] The at least one temperature sensor located downstream of the heat exchanger may be located in the fuel return line.

[0282] The method may include modulating the flow of fuel along the fuel return line using the modulating valve, based at least in part on temperature data from the at least one temperature sensor located downstream of the heat exchanger and / or the temperature sensor located in the fuel tank.

[0283] In the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth and fourteenth aspects:

[0284] The temperature of the fuel in the fuel tank may be between -54°C and 100°C.

[0285] The temperature of the fuel in the fuel tank can be between -54°C and 65°C.

[0286] The temperature of the fuel in the fuel tank may be between -54°C and 55°C.

[0287] The temperature of the fuel in the fuel tank may be between -54°C and 5°C.

[0288] The temperature of the fuel in the fuel tank may be -54°C, -50°C, -45°C, -40°C, -35°C, -30°C, -25°C, -20°C, -15°C, -10°C, -5°C, 0°C, 5°C, 10°C, 15°C, 20°C, 25°C, 30°C, 35°C, 40°C, 45°C, 50°C, 55°C, 60°C, 65°C, 70°C, 75°C, 80°C, 85°C, 90°C, 95°C, 100°C or any defined range between any two of these values.

[0289] In the first, third, fifth, seventh, ninth, eleventh, thirteenth aspects:

[0290] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based, at least in part, on an amount of fuel remaining in the fuel tank. The amount of fuel remaining in the fuel tank may be a mass of fuel remaining in the fuel tank.

[0291] The modulating valve may be adapted to modulate the flow of fuel along the fuel return line based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the fuel tank.

[0292] The modulating valve may be configured to modulate the flow of fuel along the fuel return line such that an equilibrium temperature of the fuel in the fuel tank after the fuel is returned to the fuel tank along the fuel return line does not exceed a predetermined upper threshold temperature for the fuel in the fuel tank.

[0293] In the second, fourth, sixth, eighth, tenth, twelfth, fourteenth aspects:

[0294] The method may include modulating the flow of fuel along the fuel return line using the modulating valve in operation, at at least in part, of an amount of fuel remaining in the fuel tank. The amount of fuel remaining in the fuel tank may be a mass of fuel remaining in the fuel tank.

[0295] The method may include modulating the flow of fuel along the fuel return line using the modulating valve based, at least in part, on a temperature of the fuel that has passed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0296] The method may include modulating the flow of fuel along the fuel return line using the modulating valve such that an equilibrium temperature of the fuel in the fuel tank after the fuel is returned to the fuel tank along the fuel return line does not exceed a predetermined upper threshold temperature for the fuel in the fuel tank.

[0297] As indicated elsewhere herein, the present disclosure may apply to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a bypass gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop engine, or a turbojet engine. Any such engine may or may not be provided with an afterburner. Such a gas turbine engine may be, for example, designed for land-based or marine power generation applications.

[0298] A gas turbine engine in accordance with any aspect of the present disclosure may include an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propeller).

[0299] Where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two counter-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propellers may rotate in opposite directions such that one rotates clockwise and the other counter-clockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage arranged downstream of the propeller stage. The guide vane stage may be variable pitch. Thus, high-pressure, intermediate-pressure, and free-power turbines can drive high-pressure and intermediate-pressure propellers and compressors respectively through suitable interconnecting shafts. Thus, the propellers can provide the majority of the propulsion thrust.

[0300] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reduction gear may be of the type described herein.

[0301] An engine according to the present disclosure may be a bypass engine. Such an engine may be a geared bypass engine in which the fan is directly connected to the fan drive turbine, for example without a gearbox, via a core shaft. In such a direct-drive bypass engine, the fan may be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft may be intended to rotate at a higher rotational speed than the first core shaft.In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0302] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan that is driven via a gear reducer. Accordingly, such a gas turbine engine may include a gear reducer that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gear reducer may be directly from the core shaft, or indirectly from the core shaft, for example via an intermediate shaft and / or spur gear. The core shaft may interlock the turbine and compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0303] 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 that connect turbines and compressors, e.g., one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor and the second core shaft may be intended to rotate at a higher rotational speed than that of the first core shaft.

[0304] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be adapted to receive (e.g., directly receive, e.g., via a generally annular conduit) a flow from the first compressor.

[0305] The reducer may be intended to be driven by the core shaft that is designed to rotate (e.g., in use) at the lowest rotational speed (e.g., the first core shaft in the example above). For example, the reducer may be intended to be driven only by the core shaft that is designed to rotate (e.g., in use) at the lowest rotational speed (e.g., only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the reducer may be intended to be driven by any one or both of the shafts, e.g., the first and / or second shafts in the example above.

[0306] The reducer may be a reduction gearbox (in that the output to the fan has a lower rotational speed than the input from the core shaft). Any type of reducer may be used. For example, the reducer may be a "planetary" or "star" reducer, as described in more detail elsewhere in this document. Such a reducer may be single-stage. Alternatively, such a reducer may be a compound reducer, for example a compound planetary reducer (which may have the input on the sun pinion and the output on the ring gear, and thus be referred to as a "compound star" reducer), for example with two reduction stages.

[0307] The reducer may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example in the range 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 reduction ratio may be, for example, between any two of the values ​​in the preceding sentence. Strictly by way of example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.1 or 3.2 to 3.8. Strictly by way of further example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.0 to 3.1. Strictly by way of example, the reducer may be a "planetary" reducer having a reduction ratio in the range of 3.6 to 4.2.In some arrangements the reduction ratio may be outside these ranges.

[0308] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustor, which may be provided downstream of the fan and the compressor(s) with respect to the flow path (e.g., axially downstream). For example, the combustor may be directly downstream of (e.g., at the outlet of) the second compressor, when a second compressor is provided. As a further example, the flow at the outlet to the combustor may be supplied at the inlet of the second turbine, when a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0309] The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. For example, the gas turbine engine may be a direct-drive, bypass gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first (or "low pressure") compressor and either 10 or 11 stages in the second (or "high pressure") compressor.As a further example, the gas turbine engine may be a geared gas turbine engine (in which the fan is driven by a first core shaft via a reduction gearbox) having 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may have 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As a further example, the gas turbine engine may be a geared gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0310] The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, or vice versa, as required. The respective rows of rotor blades and stator vanes may be axially offset from each other. The second (or "high pressure") turbine may comprise 2 stages in any arrangement (e.g., regardless of whether it is a geared or direct-drive engine). The gas turbine engine may be a direct-drive gas turbine engine comprising a first (or "low pressure") turbine having 5, 6, or 7 stages. Alternatively, the gas turbine engine may be a geared gas turbine engine comprising a first (or “low pressure”) turbine having 3 or 4 stages.

[0311] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or a 0% span position, to a tip at a 100% span position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or in the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, within the range of 0.28 to 0.32, or 0.29 to 0.30.These ratios can be commonly referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward-most) portion 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 outboard of any platform.

[0312] The fan radius can be measured between the engine centerline and the tip of a fan blade at its leading edge. The blower diameter (which may simply be twice the blower radius) may be greater than (or of the order of) any of: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 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 may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. Strictly by way of non-limiting example, the blower diameter may be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0313] The rotation speed of the blower may vary during use. Generally, the rotation speed is lower for blowers with a larger diameter. Strictly by way of non-limiting example, the fan rotational speed under cruising conditions may be less than 3500 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Strictly by way of further non-limiting example, the fan rotational speed under cruising conditions for a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm. Strictly by way of further non-limiting example, the fan rotational speed under cruising conditions for a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 2500 to 2800 rpm.Strictly by way of further non-limiting example, the fan rotational speed under cruising conditions for a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. Strictly by way of further non-limiting example, the fan rotational speed under cruising conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3600 to 3900 rpm. Strictly by way of further non-limiting example, the fan rotational speed under cruising conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

[0314] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a displacement of the fan blade tip with a velocity Utip. The work done by the fan blades 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 1-D average enthalpy increase) 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 peak fan load at cruise conditions may be greater than (or of the order of) any one of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values ​​being dimensionless). The peak fan load may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, within . the range from 0.28 to 0.31 or from 0.29 to 0.3 (e.g. for a geared gas turbine engine).

[0315] Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements the cruise bypass ratio may be greater than (or of the order of) any of the following: 9. 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The cruise bypass ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, within the range 12 to 16, 13 to 15, or 13 to 14.Strictly by way of non-limiting example, the cruise bypass ratio of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. Strictly by way of further non-limiting example, the cruise bypass ratio of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be at least substantially annular. The bypass duct may be radially outboard of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0316] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before an inlet into the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein under cruise conditions may be greater than (or in the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 50 to 70.Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. Strictly at . As a non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. Strictly as a non-limiting example, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0317] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, specific thrust may depend, for a given thrust condition, on the specific fuel composition supplied to the combustion chamber. At cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or in the order of) any of the following: 110 Nkg 's, 105 Nkg 's, 100 Nkg 's, 95 Nkg 's, 90 Nkg 's, 85 Nkg 's or 80 Nkg 's. The specific thrust may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example in the range of 80 Nkg 's to 100 Nkg 's, or of 85 Nkg 1 s to 95 Nkg 's. Such engines may be particularly efficient compared to conventional gas turbine engines.Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg 's to 95 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg 's to 90 Nkg 1 s. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg 's to 90 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 90 Nkg's to 120 Nkg's.

[0318] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN.Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN. The thrust referred to above may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure of 101.3 kPa, temperature of 30 degrees C), with the engine static.

[0319] In use, the temperature of the flow at the inlet of the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself may be referred to as the nozzle guide vane. In some examples, the TET may depend, for a given thrust condition, on the specific composition of fuel supplied to the combustion chamber. Under cruise conditions, the TET may be at least (or in the order of) any one of the following values: 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K, or 1650 K. The TET under cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 1530 K to 1600 K.Thus, solely by way of non-limiting example, the cruise TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1540 K to 1600 K. Strictly by way of non-limiting example, the cruise TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the cruise TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1600 K to 1660 K. Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1570 K to 1630 K.

[0320] The maximum TET under engine use may be, for example, at least (or in the order of) any of the following values: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. The maximum TET may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range from 1800 K to 1950 K, or from 1900 K to 2000 K.Thus, solely by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. blower in the range of 300 cm to 340 cm can be in the range of 1935 K to 1995 K.Strictly by way of non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum takeoff thrust (PMD) condition.

[0321] A fan blade and / or airfoil portion of a fan blade described and / or claimed herein may be fabricated from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part 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 least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to withstand impact (e.g., by birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade may have a carbon fiber-based or aluminum-based (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0322] A blower as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be connected to the central portion in any desired manner. For example, each fan blade may comprise a fastening element that may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail that may notch into and / or engage a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a finned disc or finned ring.Any suitable method may be used to manufacture such a bladed disc or finned ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be joined to the hub / disc by welding, such as a linear friction stir weld.

[0323] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the bypass duct exit area to be varied during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0324] The fan of a gas turbine engine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. When the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. When the fan blades have a metal body (for example, aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0325] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach and landing (or one or more parts thereof) have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, one skilled in the art would immediately recognize that each term refers to all, or one or more parts, of a phase of operation of the engine within a given mission of an aircraft to which the gas turbine engine is designed to be attached.

[0326] As such, ground idling may refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a need for the engine to taxi. At idle, the engine may produce between 3% and 9% of the available engine thrust. In other non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. When taxiing, the engine may produce between 5% and 15% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In other non-limiting examples, the engine may produce between 7% and 10% of the available thrust.Takeoff may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In other non-limiting examples, the engine may produce 100% of the available thrust.

[0327] Climb may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In other non-limiting examples, the engine may produce between 85% and 90% of the available thrust. As such, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and arrival at cruise conditions, with arrival at cruise conditions thus defining the start of the cruise phase, or a portion thereof, of the aircraft's flight.Additionally or alternatively, climb may refer to a nominal point in, or one or more nominal periods during, an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may require an additional demand for engine thrust.

[0328] As used herein, cruise conditions, which may define the cruise phase (or a portion thereof) of aircraft flight, have the conventional meaning and will be readily understood by those skilled in the art. In some examples, for a given gas turbine engine for an aircraft, cruise conditions may refer to the mid-cruise engine operating point of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. As such, mid-cruise may be considered the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the beginning of descent has been burned (which may be approximated by the midpoint - in terms of time and / or distance - between the end of climb and the beginning of descent).Cruise conditions may therefore define an operating point, phase, or part thereof, of flight that provides thrust to ensure steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number), or at least substantially steady-state operation (i.e., maintaining at least a substantially constant altitude and / or at least a substantially constant Mach number) of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine may provide half the total thrust that would be required for steady-state operation, or at least substantially steady-state operation, of that aircraft at mid-cruise.

[0329] In other words, for a given gas turbine engine for an aircraft, cruise conditions can be defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady-state operation, or at least substantially steady-state operation, of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine at cruise conditions can be clearly defined.

[0330] Strictly by way of example, the forward speed in cruise condition may be any point in the range of Mach 0.7 to 0.9, e.g., 0.75 to 0.85, e.g., 0.76 to 0.84, e.g., 0.77 to 0.83, e.g., 0.78 to 0.82, e.g., 0.79 to 0.81, e.g., of the order of Mach 0.8, of 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 condition. For a given aircraft, cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

[0331] Strictly by way of example, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

[0332] Strictly by way of example, cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of 20 kN to 40 kN.

[0333] Strictly by way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is of course dependent on the engine and its intended application and may be, for example, a value ranging from 35 kN to 65 kN.

[0334] In use, a gas turbine engine described and / or claimed herein may operate at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0335] Further, one skilled in the art would readily recognize that either or both of a descent and an approach refer to a phase of operation within an aircraft flight cycle between cruise and landing. of the aircraft, with the approach in particular forming part of the landing and takeoff (LTO) phase. During either or both of the descent and the approach, the engine may produce between 0% and 50% of the available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of the available thrust. In other non-limiting examples, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing, where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

[0336] According to one aspect, there is provided an aircraft 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. Accordingly, the cruise conditions according to this aspect may correspond to an operating point, a phase, or a portion thereof, of the flight of the aircraft, as defined elsewhere herein.

[0337] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be at any suitable condition, which may be as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).

[0338] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may comprise (or may be) operation at any suitable condition, for example at mid-cruise of the aircraft, as defined elsewhere herein.

[0339] Those skilled in the art would understand that, unless mutually exclusive, any feature or parameter described in connection with any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter contained or described herein may be applied to any aspect and / or in association with any other feature or parameter contained or described herein.

[0340] Unless mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with any other parameter(s) and / or value(s) contained or described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with any other parameter(s) and / or value(s) contained or described herein (e.g., parameter B; parameter C; and parameter D, and so on) to express a product of their relationship. For example, those skilled in the art will understand that when parameter A is described separately from parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or any other application, combination or function of parameter A in relation to parameter B, as required.

[0341] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0342] [Fig.l] is a sectional side view of a gas turbine engine;

[0343] [Fig.2] is a close-up sectional side view of an upstream portion of a motor geared gas turbine;

[0344] [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine;

[0345] [Fig.4] is a close-up sectional side view of an upstream portion of a direct-drive gas turbine engine;

[0346] [Fig.5] is a schematic representation of an exemplary fuel system including a fuel return line;

[0347] [Fig.6] is a schematic representation of an exemplary fuel system, including a fuel recirculation line;

[0348] [Fig.7] is a schematic representation of another example fuel system including a fuel return line;

[0349] Figures 8 to 13 illustrate examples of methods of operating a gas turbine engine;

[0350] [Fig. 14] is a schematic representation of an exemplary fuel system, including a fuel recirculation line;

[0351] [Fig. 15] is a schematic representation of an exemplary fuel system, including a fuel recirculation line and a fuel return line;

[0352] [Fig. 16] is a schematic representation of an exemplary fuel system, including a fuel recirculation line; and

[0353] [Fig. 17] is a representation of an aircraft with a propulsion system comprising two gas turbine engines.

[0354] [Fig.l] illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 comprises an air intake 12 and a propulsion fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 which receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a 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 reduction gear 30.

[0355] In use, the core air stream A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustor 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby 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 by a suitable interconnecting shaft 27.The blower 23 generally acts to impart increased pressure to the bypass airflow B flowing through the bypass duct 22, such that the bypass airflow B is exhausted through the bypass exhaust nozzle 18 to generally provide the majority of the propulsion thrust. The epicyclic reduction gear 30 is a reduction box.

[0356] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in [Fig. 2]. The low pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun gear, or sun pinion, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and meshing therewith are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation about the sun gear 28 in synchronism while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause it to rotate around the engine axis 9.Radially outwardly of and meshing with the planet gears 32 is a ring or crown gear 38 which is coupled, via links 40, to a stationary support structure 24.

[0357] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the reduction gear output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as an "intermediate pressure turbine" and "intermediate pressure compressor." When such alternative nomenclature is used, the fan 23 may be designated as the first compression stage or the lower pressure compression stage.

[0358] The epicyclic reduction gear 30 is shown by way of example in more detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic planetary reduction gear 30 generally include at least three planet gears 32.

[0359] The epicyclic reduction gear 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reduction gear 30 may be used. As a further example, the epicyclic reduction gear 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the reduction gear 30 may be a differential reduction gear in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0360] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the reducer 30 to the motor 10. As a further example, the connections (such as the links 36, 40 in the example of [Fig. 2]) between the reducer 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of rigidity or flexibility.As a further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g., between input and output shafts from the reducer and stationary structures, such as the reducer housing) may be used, and the description is not limited to the exemplary arrangement of [Fig. 2]. For example, when the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that . the arrangement of output and support links and bearing locations would typically be different from that shown as an example in [Fig.2].

[0361] Thus, the present disclosure extends to a gas turbine engine having any arrangement of reduction gear styles (e.g., star or planet gear), support structures, input and output shaft arrangements, and bearing locations.

[0362] Optionally, the reducer may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster).

[0363] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have a different number of compressors and / or turbines and / or a different number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.l] has a split-flow nozzle 18, 20 which means that the flow through the bypass duct 22 has its own nozzle 18 which is independent of, and radially outward of, the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied 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 referred to as a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) may have a fixed or variable area.

[0364] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a reduction gear for the main shaft(s), being instead direct drive engines. A sectional view of such an engine is shown in [Fig.4].

[0365] With reference to [Fig.4], a gas turbine engine is generally indicated at 10, having a main axis of rotation 9. The engine 10 comprises, in axial flow series, an air intake 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0366] In use, air entering the intake 12 is accelerated by the blower 23 to produce two airflows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate pressure compressor 14, and the bypass airflow B passes through a bypass duct 22 to provide propulsive thrust. The pressure compressor intermediate 14 compresses the air flow A before delivering this air to the high pressure compressor 15 where further compression takes place.

[0367] Compressed air discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustion chamber 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive, the high-pressure, intermediate-pressure, and low-pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional propulsive thrust. The high-pressure 17, intermediate-pressure 19a, and low-pressure 19 turbines respectively drive the high-pressure compressor 15, the intermediate-pressure compressor 14, and the fan 23, each by a suitable interconnecting shaft.

[0368] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a reduction gear provided in the drive train from a turbine to a compressor and / or a fan.

[0369] While the described example relates to a bypass engine, the description may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[0370] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-up direction in [Fig.l]) and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial and circumferential directions are mutually perpendicular.

[0371] The fuel F supplied to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may comprise molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. Additionally or alternatively, when blended with, mixed with, or substituted for an alternative fuel, the fuel F may comprise renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F may comprise one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganic substances, and metals.

[0372] By SAF, the skilled person means, for example, a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel or a biofuel, produced from biological or non-biological resources. It is understood that SAFs are commonly synthesized from carbon-containing gases extracted from the atmosphere and / or captured during industrial processes; or from a wide range of sustainable feedstocks such as, for example, used oils and greases; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata and pennycress; non-biogenic alternative fuels; jatropha; halophytes and algae, rather than fossil-based hydrocarbons. It is understood that SAFs do not include fossil fuels.

[0373] The functional performance of a given fuel composition, or fuel mixture F for use in a given mission, may be defined, at least in part, by the fuel's ability to process the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and emissions comprising gaseous and / or particulate matter. As such, particulate matter emissions may include soot particles created by the combustion of said fuel F, also known as non-volatile particulate matter (nvPM). Any reference to soot or smoke herein may also apply to other types of particulate matter emissions known in the art.The gaseous emissions may include one or more nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO), including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOCs) created by the combustion of said fuel F. Any reference to gaseous emissions herein may also apply to other types of gaseous emissions known in the art.

[0374] A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, may at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, may at least partially reduce takeoff fuel volume, which may be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibiting degradation or coking of fuel under thermal stress) can allow the fuel to experience elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency. Reduced emissions, including particulate matter, can allow for reduced contrail formation, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance.For example, a relatively lower freezing point (°C) may allow long-range missions to optimize flight profiles; minimum aromatic concentrations (%) may ensure sufficient swelling of some materials used in the construction of O-rings and seals previously exposed to high aromatic fuels; and, a maximum surface tension (mN / m) may ensure sufficient spray breakup and atomization of the fuel.

[0375] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels having higher ratios of hydrogen atoms to carbon atoms may have higher specific energies in the absence of bonding strain. For example, fossil-based hydrocarbon fuels may comprise molecules having about 7 to 18 carbons, with a significant portion of a given composition derived from molecules of 9 to 15 carbons, with an average of 12 carbons.

[0376] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with further compositions pending approval. However, there is an anticipation in the aviation industry that sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) will eventually be approved for use. Any reference to "SAF" herein may refer to a fuel composed of 100% SAF or to a fuel comprising SAF, e.g., a blend of SAF.

[0377] Sustainable aviation fuels may include one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may include aromatic contents and / or lower sulfur content, relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may include one or both of a higher isoalkane content and a higher cycloalkane content, relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may include a density of between 90% and 98% of that of kerosene and / or a calorific value of between 101% and 105% of that of kerosene.

[0378] In some examples, the sustainable aviation fuel(s) or blend(s) supplied to the combustion equipment 16 may have a relatively lower aromatics and / or other non-paraffinic content than kerosene. The sustainable aviation fuel may have an aromatics content of e.g., 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; e.g., 4%, 3%, 2%, 1%, or less than 1%; e.g., 0.75%, 0.5%, 0.25%, or less than 0.25%; e.g., 0.2%, 0.1%, or less than 0.1%; e.g., 0.01%, 0.001%, or 0%. The aromatics content of sustainable aviation fuel may be within one figure or a range bounded by any two or more of the values ​​in the preceding sentence (i.e. the values ​​may form upper or lower limits), e.g.13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or from 0% to 0.75%, from 0% to 0.5%, or from 0.1% to 0.25%; or from 0.15% to 0.65%, from 0.35% to 0.55%, or from 0.035% to 0.055%; depending on one or more preferences, fuel stock or supplier, and composition variations.

[0379] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide beneficial effects including, for example, one or more of higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; less NOx; and, lower CO2 emissions, compared to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment 16).Thus, compared to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to either or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs.

[0380] An exemplary fuel system 1000 comprising a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 is shown schematically in [Fig. 5]. In the illustrated example, the engine 10 comprises a reduction gear 30, as shown in Figures 1-3 (it can, however, be used with other gas turbine engines, including those with a direct-drive architecture). The fuel system 1000 includes both the fuel supply system (including the fuel tank 50 and the pump 1002), which supplies fuel to the engine, and the fuel management system 1500 (which operates within the engine) of the aircraft. The fuel management system 1500 manages the fuel temperature as well as the fuel flow, directing the fuel through one or more heat exchangers 1004, 1006 of the engine heat exchange system.

[0381] In the described implementation, each engine 10 includes its own fuel management system 1500. In other implementations, a single fuel management system 1500 may manage the fuel supply to multiple engines, and may, for example, include a duplicate of various of the elements shown in [Fig. 5] for the other engine(s).

[0382] Fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low-pressure fuel feed pump 1002. The fuel then passes through a secondary fuel-oil heat exchanger 1004 and a primary fuel-oil heat exchanger 1006 before passing through an engine fuel pump 1003. The engine fuel pump 1003 may be described as a primary or high-pressure fuel pump. The primary fuel-oil heat exchanger 1006 may be referred to as a primary fuel-oil heat exchanger because the oil flowing through it may be used to cool and lubricate the main reduction gear 30 of the engine 10.The secondary fuel-oil heat exchanger 1004 may be described as a generator fuel-oil heat exchanger, as oil flowing therethrough may be used to cool and / or lubricate a generator of the engine 10 designed to provide aircraft electrical power (e.g., an integrated drive generator (IDG) of the engine 10). The illustrated fuel management system 1500 is intended to cause fuel to reach the secondary fuel-oil heat exchanger 1004 before the primary fuel-oil heat exchanger 1006.

[0383] The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are designed such that an oil flow is routed through each in addition to the fuel flow therethrough. The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are designed such that heat can be transferred between the oil and the fuel flowing therethrough. Under cruising conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006, and the average temperature of the oil flow entering the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1004. In this manner, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are each configured to transfer thermal energy from an oil flow to a fuel flow flowing therethrough under cruising conditions. The fuel system 1000 includes an electronic controller configured to control the operation of the heat exchangers 1004, 1006.

[0384] The primary fuel-oil heat exchanger 1006 may have oil used to lubricate and / or cool a main reduction gear 30 of the gas turbine engine 10 passing therethrough, and may thus be described as a primary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1004 may have oil used to lubricate and / or cool one or more components of a generator of the engine 10 passing therethrough, and may thus be described as a generator fuel-oil heat exchanger.

[0385] The two oil flows may be physically separated from each other and possibly also chemically distinct and / or having a different flow rate. A different oil than that flowing through the secondary fuel-oil heat exchanger 1004 may therefore flow through the primary fuel-oil heat exchanger 1006.

[0386] Generally, at least the majority of the fuel passing through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006, although either or each of the heat exchangers 1004, 1006 may be provided with a bypass to allow some of the fuel to avoid passing through the respective heat exchanger, for example in the form of a bypass pipe 1005 as illustrated in [Fig. 5]. A valve (not shown) may determine what proportion of the fuel passes through the heat exchanger 1004 and what proportion passes through the bypass pipe 1005.

[0387] The two heat exchangers 1004, 1006 may therefore be described as being in series with each other, with respect to the fuel flow. In general, at least the majority of the fuel passing through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006, although each heat exchanger 1004, 1006 may be provided with a bypass to allow some of the fuel to avoid passing through the respective heat exchanger, for example in the form of a bypass pipe 1005. A valve may determine what proportion of the fuel passes through the heat exchanger 1004 and what proportion passes through the bypass pipe 1005. Additionally or alternatively, one or more bypass pipes may also be provided to allow oil to bypass one or more heat exchangers.

[0388] The secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006 are designed such that an oil flow is also carried through each of them. The oil that flows through one fuel-oil heat exchanger is different from the oil that flows through the other fuel-oil heat exchanger in the described implementation, although it is obvious that the same oil may flow through one fuel-oil heat exchanger and then through another fuel-oil heat exchanger in other implementations. In such implementations, one or more bypass pipes for oil or fuel may again be provided to allow the fluid(s) to bypass one or more heat exchangers.

[0389] The two heat exchangers 1004, 1006 are thus in closed loop systems separate with respect to the oil flow, in the implementation being described, namely the oils flowing through the primary and secondary fuel-oil heat exchangers are fluidically separated, and may be chemically distinct from each other.

[0390] The fuel system 1000 further includes a modulator valve 1010 located downstream of the primary fuel-oil heat exchanger 1006 and configured to divert or direct at least a portion of the fuel that has left the primary fuel-oil heat exchanger back to the fuel tank 50 via a fuel return line 1011. The modulator valve 1010 is configured to modulate a flow of fuel along the fuel return line 1011. The modulator valve 1010 can determine what proportion of the fuel is returned to the fuel tank 50, and what proportion continues to the combustion chamber 16. It will be appreciated that the modulator valve 1010 can be located at substantially any position downstream of at least one of the heat exchangers 1004, 1006.For example, the modulating valve 1010 may alternatively be located downstream of the secondary heat exchanger 1004 and upstream of the primary heat exchanger 1006. It will also be appreciated that the fuel system 1000 may alternatively include a single fuel-oil heat exchanger or may include more than two fuel-oil heat exchangers, and that the modulating valve 1010 may be located downstream of at least one of the heat exchangers (either directly downstream or with one or more intermediate components of the fuel system 1000 between the modulating valve 1010 and the heat exchanger).

[0391] The heat exchangers illustrated in [Fig.5] and described herein are examples only. The primary and / or secondary heat exchangers may provide cooling for other aircraft systems than those described above.

[0392] In the illustrated example, the modulating valve 1010 is located downstream of the primary fuel-oil heat exchanger 1006. In the illustrated example, the modulating valve 1010 is positioned upstream of the engine fuel pump 1003, although this is not essential. The modulating valve 1010 is intended to allow a controlled quantity of fuel to be returned to the fuel tank 50.

[0393] The fuel system 1000 includes a plurality of temperature sensors (indicated by the circular symbols surrounding the capital letter T) configured to measure a temperature of the fuel at different locations in the fuel system 1000. In the illustrated example, the fuel system 1000 includes a temperature sensor located in the fuel tank 50 for measuring a temperature of the fuel in the fuel tank 50. Additionally or alternatively, the fuel system 1000 may include a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006, which may be used to determine or be representative of a temperature of fuel in the fuel tank 50.

[0394] In the illustrated example, the fuel system 1000 includes two temperature sensors located downstream of the heat exchangers 1004, 1006. One of the temperature sensors is located in the fuel return line 1011, while the other temperature sensor is located downstream of the modulating valve 1010 and upstream of the combustion chamber 16. Either or both of the temperature sensors may be used to determine or represent a temperature of the fuel delivered to the combustion chamber 16 and / or a temperature of the fuel returned to the fuel tank 50. It will be appreciated that only one of the temperature sensors located downstream of the heat exchangers 1004, 1006 may be provided.Depending on the position of the modulating valve 1010 relative to one or more heat exchangers (e.g., the primary heat exchanger 1006 and / or the secondary heat exchanger 1004), it will be appreciated that the temperature sensor(s) may be placed at any suitable location downstream of the heat exchanger(s) in order to measure temperature data (or provide representative temperature data) relating to a temperature of the fuel delivered to the combustion chamber 16 and / or a temperature of the fuel returned to the fuel tank 50.

[0395] Returning fuel to the fuel tank 50 provides a mechanism for controlling fuel flow from the fuel system 1000, for example to manage thermal loads in the engine 10 and / or to control a fuel temperature at different locations in the system 1000 (such as in the fuel tank 50 or at the inlet of the combustion chamber 16).

[0396] The modulating valve 1010 may be adapted to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 in any suitable manner. The fuel system 1000 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulating valve 1010.

[0397] In one example, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of a temperature, in kelvin, of the fuel in the fuel tank 50 to a temperature, in kelvin, of the fuel delivered to the combustion chamber 16 under cruising conditions is less than 0.56. The ratio of the temperature of the fuel in the fuel tank 50 to the temperature of the fuel delivered to the combustion chamber 16 may also be referred to or referred to as the AT ratio (or simply AT).A fuel temperature measured downstream of the heat exchangers 1004, 1006 (e.g., measured by a temperature sensor located downstream of one or both of the heat exchangers 1004, 1006) may be used to determine or be representative of a temperature of the fuel delivered to the combustion chamber 16 (i.e., the fuel temperature at the inlet of the combustion chamber). The fuel temperature in the fuel tank 50 may be measured using a temperature sensor located in the fuel tank 50 or a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006. The electronic fuel system controller 1000 is configured to control operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0398] The AT ratio is calculated using the temperature of the fuel in the fuel tank 50 in kelvin (K) and the temperature of the fuel delivered to the combustion chamber 16 in kelvin (K), 0°C being equal to 273.15 K.

[0399] A temperature of the fuel in the fuel tank 50 is typically maintained between -54°C (219.15 K) and 65°C (338.15 K). The lower limit of -54°C is typically used to prevent the fuel from freezing in the fuel tank 50. Alternatively, the upper limit of the temperature of the fuel in the fuel tank 50 may be 55°C (328.15 K), or 5°C (278.15 K), or 0°C (273.15 K). In some other examples, the upper limit of the temperature of the fuel in the fuel tank 50 may be 100°C (373.15 K).

[0400] The maximum fuel temperature in the fuel tank may be controlled by the flash point of the fuel to reduce the generation of fuel vapors that could ignite in the presence of an ignition source. When fuel is not returned to the fuel tank, from an engine operating perspective, it may be advantageous for the fuel in the tank to be cooler, but not so much that it causes a fuel icing problem in the engine fuel system. If the fuel tank temperature is low, more of the heat generated within the engine can be returned to the fuel, which is more efficient.The inventors have found that when fuel return to the tank is used, the system limits extend from the engine alone to the aircraft engine and fuel tank combined, so that a cooler tank is not necessarily more advantageous and the fuel in the tank can be maintained at a warmer temperature.

[0401] For sustainable aviation fuel (i.e., 100% S AF fuel or S AF blend), the temperature of the fuel delivered to the combustion chamber 16 under cruise conditions is typically at least 120°C (393.15 K), or typically between 120°C (393.15 K) and 180°C (453.15 K) or higher (e.g., 200°C or 473.15 K). Preferably, the temperature of the fuel delivered to the combustion chamber 16 under cruise conditions may be typically at least 140°C (413.15 K), or typically between 140°C (413.15 K) and 180°C (453.15 K) or higher (such as 200°C or 473.15 K).

[0402] Alternatively, the temperature of the fuel delivered to the combustion chamber 16 under cruising conditions may be at least 120°C, 125°C, 130°C, 135°C, 140°C, 145°C, 150°C, 155°C, 160°C, 165°C, 170°C, 175°C, 180°C, 185°C, 190°C, 195°C or 200°C, or a fuel temperature defined between any two of these values.

[0403] The temperature of the "fuel delivered to the combustion chamber", the temperature of the "fuel supplied to the combustion chamber" and the temperature of the "fuel at the inlet of the combustion chamber" are used interchangeably herein and should be considered to refer to the same temperature.

[0404] A temperature of the fuel delivered to the combustion chamber 16 under cruising conditions may be defined as an average for at least 5 minutes, and possibly for ten minutes, under steady-state cruising conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in operating fuel temperature, often a rise in temperature. Each fluctuation may last no longer than 5 minutes. The electronic control device of the system fuel 1000 is configured to control the operation of the heat exchangers 1004, 1006 to control a temperature of the fuel delivered to the combustion chamber 16.

[0405] The modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is less than 0.56. This value corresponds to less than -54°C / 120°C. The inventors have identified that this range can be achieved using SAFs where the fuel temperature at the combustion chamber inlet is not limited to 120°C. This is not the case for a fossil kerosene-based fuel due to the risk of thermal degradation of the fuel. The minimum fuel temperature in the fuel tank is limited to -54°C in this example to avoid the risk of freezing. The value of AT in this example (and in other similar examples where only an upper limit is given) would be greater than zero.

[0406] Generally speaking, heating the fuel to a higher temperature upon delivery to the combustion chamber can improve the thermodynamic efficiency of the engine. The inventors have determined that the use of SAF can allow for a higher fuel temperature at the combustion chamber inlet without risking significant thermal degradation of the fuel that might otherwise occur if fossil kerosene-based fuel were used. The inventors have determined how the engine's fuel / heat management system can be advantageously controlled to utilize this difference between SAF and kerosene. Increasing the fuel temperature can also be beneficial by allowing more heat to be transferred from the oil within the heat exchanger(s) to improve cooling.

[0407] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is less than 0.53. This value corresponds to less than -54°C / 140°C. This corresponds to a minimum temperature of 140°C for the fuel supplied to the combustion chamber and the minimum temperature of -54°C for the fuel in the fuel tank. The inventors have determined that it is possible to operate in this regime using the SAF, as it can be advantageously heated to a higher temperature to improve thermodynamic efficiency without the risk of thermal degradation.

[0408] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.48 to 0.56. This corresponds to between -54°C / 180°C and -54°C / 120°C. This corresponds to a fuel temperature at the combustion chamber inlet being between 120°C and 180°C, while the minimum fuel temperature in the fuel tank remains -54°C. The maximum limit of 180°C may be beneficial in reducing the risk of fuel degradation that can occur even if SAF is used.

[0409] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.48 to 0.53. This corresponds to between -54°C / 180°C and -54°C / 140°C. This corresponds to a fuel temperature at the inlet of the combustion chamber being between 140°C and 180°C, while the minimum fuel temperature in the fuel tank remains -54°C. This is again a better use of the thermal properties of the SAFs to improve thermodynamic efficiency without the risk of fuel degradation at too high a temperature.

[0410] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.48 to 1.00. This corresponds to between -54°C / 180°C and 65°C / 65°C. The lower limit corresponds to a fuel temperature at the combustion chamber inlet equal to the upper limit of 180°C mentioned above, while the minimum fuel temperature in the tank is limited to -54°C to reduce the risk of freezing. The upper limit is 1.00 because the heat exchanger(s) of the present examples are not designed to operate in a condition where they remove heat from the fuel, i.e., the fuel will not be colder at the combustion chamber inlet than its temperature in the fuel tank.

[0411] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.48 to 0.90. This corresponds to between -54°C / 180°C and 100°C / 140°C. This upper lower limit of 0.90 compared to the example in the previous paragraph corresponds to a maximum fuel temperature in the fuel tank of 100°C, as indicated above, with the minimum fuel temperature at the combustion chamber inlet being 140°C, in order to utilize the superior thermal properties of the SAF.

[0412] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.48 to 0.82. This corresponds to between -54°C / 180°C and 65°C / 140°C. The upper limit of 0.82 corresponds to a maximum fuel temperature in the fuel tank of 65°C, as noted above, with the minimum fuel temperature at the inlet of the combustion chamber being 140°C in order to utilize the superior thermal properties of SAF to improve thermodynamic efficiency.

[0413] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.56 to 1.00. This corresponds to between -54°C / 140°C and 65°C / 65°C. The upper limit of 1.00 corresponds to the maximum that can be achieved without the fuel being hotter at the fuel tank than at the combustion chamber. The lower limit corresponds to a fuel temperature equal to or greater than 140°C at the inlet of the combustion chamber, in order to utilize the thermal properties of the SAF to improve thermodynamic efficiency, while the minimum temperature at the fuel tank is -54°C to reduce the risk of freezing.

[0414] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.56 to 0.82. This corresponds to between -54°C / 140°C and 65°C / 140°C. This corresponds to the same lower limit as in the example in the previous paragraph. The upper limit of 0.82 corresponds to a maximum fuel temperature in the fuel tank of 65°C, as noted above, with the minimum fuel temperature at the combustion chamber inlet being 140°C in order to utilize the superior thermal properties of the SAF to improve thermodynamic efficiency.

[0415] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that AT is in the range of 0.56 to 0.75. This corresponds to between -54°C / 140°C and 65°C / 180°C. This corresponds to the same lower limit as in the example in the previous paragraph. The upper limit of 0.75 corresponds to a maximum fuel temperature in the fuel tank of 65°C, as indicated above, and the maximum value of 180°C of the fuel at the inlet of the combustion chamber, as indicated above. The maximum value at the inlet of the combustion chamber can only be used in some examples when the fuel temperature in the fuel tank is relatively high in order to avoid a large temperature difference between the two.

[0416] [Fig.8] illustrates a method 2000 of operating a gas turbine engine 10. The method 2000 includes modulating 2001 the flow of fuel along the fuel return line 1011 using the modulating valve 1010 such that a ratio between a temperature of the fuel in the fuel tank and a temperature of the fuel delivered to the combustion chamber (AT) is as defined in any of the examples described above with respect to fuel system 1000 or as defined elsewhere herein.

[0417] In another example, the modulator valve 1010 is further or alternatively to modulate fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at startup of the engine 10 is less than 0.56. The electronic fuel system controller 1000 is configured to control operation of the modulator valve 1010 accordingly to achieve this value or any other value of AT at cruise to AT at startup defined herein.

[0418] The operation of the modulating valve such that the ratio between AT at cruising speed and AT at engine start 10 is less than 0.56 corresponds to a value less than ((-54°C / 120°C) / 1.00). It therefore corresponds to the values ​​that can be achieved using fuel temperatures at the combustion chamber inlet of 120°C or higher, which can be achieved by using the improved thermal properties of SAF compared to fossil fuel. The value of 0.56 corresponds to a minimum value of AT set by the minimum of -54°C for the fuel in the fuel tank to reduce the risk of freezing, with the fuel temperature at the combustion chamber inlet being at least 120°C. The maximum value of AT at start is 1, which corresponds to the maximum that can be achieved without the fuel being hotter in the fuel tank than at the combustion chamber.

[0419] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at start of the engine 10 is less than 0.53 (i.e. (-54°C / 140°C) / 1.00). This corresponds to the value of the example in the previous paragraph, except that a fuel temperature at the combustion chamber inlet is at least 140°C, which allows for further use of the thermal properties of the SAF to improve thermodynamic efficiency.

[0420] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at start of the engine 10 is between 0.48 and 0.56. This corresponds to between ((-54°C / 180°C) / l.00) and ((-54°C / 120°C) / l.00). The upper limit of 0.56 corresponds to that discussed above. The lower limit of 0.48 corresponds to a maximum combustion chamber inlet fuel temperature of 180°C, as discussed above, to reduce the risk of thermal degradation of the fuel, the minimum fuel temperature in the fuel tank being -54°C to reduce the risk of freezing.

[0421] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio between AT at cruise and AT at start of the engine 10 is between 0.48 and 0.53. This corresponds to between ((-54°C / 180°C) / l.00) and ((54°C / 140°C) / l.00). The upper limit of 0.53 corresponds to that discussed above, which makes more use of the thermal properties of the SAF. The lower limit of 0.48 corresponds to that discussed in the previous paragraph.

[0422] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at startup of the engine 10 is between 0.48 and 1.88. This corresponds to between ((-54°C / 180°C) / (100°C / 100°C)) and ((100°C / 120°C) / (-54°C / 160°C)). The inventors have determined that the limits of the ratio of AT at cruise to AT at startup can be determined by the expression:

[0423] ((min_tank_cruise / max_combustor_cruise) / (max_tank_startup / min_combus tor_startup @max tank startup)) to ((max_tank_cruise / min_ combustor _cruise) / (min_tank_startup / max_combustor _startup))

[0424] The lower limit of 0.48 corresponds to the minimum fuel temperature in the tank of -54°C to reduce the risk of freezing and the maximum fuel temperature at the combustion chamber inlet at cruising speed being 180°C to reduce the risk of thermal degradation of the fuel and improve thermodynamic efficiency. The maximum value of AT at start-up is 1 because the heat exchanger(s) are not intended to cool the fuel before it reaches the combustion chamber. The value of AT at start-up is therefore limited by the fact that the minimum fuel temperature at the combustion chamber inlet is not lower than the corresponding maximum value of the fuel in the tank for which the maximum value of AT is calculated (e.g. 100°C).The upper limit of 1.88 corresponds to a maximum fuel temperature in the tank of 100°C, a minimum fuel temperature at the combustion chamber inlet of 120°C to utilize the thermal properties of the SAF, a minimum fuel temperature in the tank of -54°C to reduce the risk of freezing, and a maximum fuel temperature at the combustion chamber inlet of 160°C during start-up.

[0425] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio between AT at cruise and AT at engine start 10 is between 0.48 and 1.71. This corresponds to between ((-54 °C / 180 °C) / (65 °C / 65 °Q) and ((65 °C / 120 °C) / (-54 °C / 160 °Q). These limits correspond to those of the previous paragraph, but with the maximum fuel temperature in the fuel tank being 65 °C.

[0426] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at startup of the engine 10 is between 0.48 and 1.70. This corresponds to between (-54°C / 180°C) / (100°C / 100°C) and (100°C / 120°C) / (-54°C / 120°C). In this example, AT at startup may be AT at startup when the aircraft is on the ground. The limits of this paragraph correspond to those of the previous paragraph, except for a maximum fuel temperature in the tank being 100°C, and the maximum fuel temperature at the combustion chamber inlet during startup being 120°C.

[0427] In some examples, the modulating valve 1010 is intended to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio between AT at cruise and AT at start of the engine 10 is between 0.48 and 1.55. This corresponds to between ((-54°C / 180°C) / (65°C / 65°Q) and ((65°C / 120°C) / (-54°C / 120°Q). In this example, AT at start may be AT at start when the aircraft is on the ground. These limits correspond to those of the previous paragraph, with the exception of a maximum temperature of the fuel in the tank being 65°C.

[0428] In some examples, the modulating valve 1010 is to modulate fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at startup of the engine 10 is between 0.48 and 1.88. This corresponds to between (-54°C / 180°C) / (100°C / 100°C) and (100°C / 120°C) / (-54°C / 160°C). In this example, AT at startup may be AT at startup when the aircraft is in flight. These limits correspond to those of the previous paragraph, except that a maximum fuel temperature in the tank is 100°C, and the maximum fuel temperature at the combustion chamber inlet during startup is 160°C.

[0429] In some examples, the modulating valve 1010 is to modulate the flow of fuel along the fuel return line 1011 to the fuel tank 50 such that a ratio of AT at cruise to AT at startup of the engine 10 is between 0.48 and 1.71. This corresponds to between (-54°C / 180°C) / (65°C / 65°C) and (65°C / 120°C) / (-54°C / 160°C). In this example, AT at startup may be AT at startup when the aircraft is in flight. These limits correspond to those of the previous paragraph, except for a maximum fuel temperature in the tank being 65°C.

[0430] In any example described herein, the fuel temperature at the combustion chamber inlet at cruise may be from 120°C to 180°C, and preferably from 140°C to 180°C. In any example described herein, the fuel temperature at the combustion chamber inlet at startup may be from -40°C to 160°C. The fuel temperature at the combustion chamber inlet at ground startup may be from -40°C to 120°C. The fuel temperature at the combustion chamber inlet at air startup may be from 0°C to 160°C.

[0431] The temperature range indicated in the preceding paragraph may correspond to a ground start where the engine starts with all components at ambient temperature. For an air start (i.e., in flight), the temperature range indicated in the preceding paragraph corresponds to a relight scenario. In this case, the engine was running and is then restarted. This may involve the aircraft descending from cruising altitude to a windmill relight envelope. In this time, the temperature of the core engine components decreases, but by the time relight occurs, there may still be some residual heat in the engine. The temperature ranges for ground restart and air restart are therefore different.

[0432] [Fig.9] illustrates a method 2100 of operating a gas turbine engine 10. The method 2100 includes modulating 2101 the fuel flow along the fuel return line 1011 using the modulating valve 1010 such that the ratio between AT at cruise and AT at start of the engine 10 is as defined in any of the examples above or defined elsewhere herein.

[0433] In another example, the modulating valve 1010 is further or alternatively to initiate a return of fuel to the fuel tank 50 when the fuel delivered to the combustion chamber 16 (or having passed through one or both of the heat exchangers 1004, 1006) is at a temperature of at least 120°C. When fuel temperatures are below 120°C, the modulating valve 1010 may prevent the return of fuel to the fuel tank 50. The electronic fuel system controller 1000 is configured to control the operation of the modulating valve 1010 accordingly. In other examples, the modulating valve may be configured to initiate a return of fuel to the fuel tank 50 when the fuel delivered to the combustion chamber 16 (or having passed through one or both of the heat exchangers 1004, 1006) is at a temperature of at least 140°C.This allows for further use of the thermal properties of the S AF. In yet other examples, the modulating valve. may be intended to initiate a return of fuel to the fuel tank when the fuel having passed through one or both heat exchangers is at a temperature between 120°C and 180°C, and preferably is at a temperature between 140°C and 180°C.

[0434] A fuel temperature is measured downstream of one or both of the heat exchangers 1004, 1006. In this example, the fuel temperature is measured using a temperature sensor in the fuel return line 1011, although it is appreciated that the fuel temperature may be measured using a temperature sensor located at substantially any position downstream of one or both of the heat exchangers 1004, 1006. A temperature of the fuel having passed through one or both of the heat exchangers 1004, 1006 may be used to determine or be representative of a temperature of the fuel returned to the fuel tank 50. The electronic fuel system controller 1000 is configured to control operation of the modulating valve 1010 based at least in part on temperature data from the one or more temperature sensors located downstream of one or both of the heat exchangers 1004, 1006.

[0435] In this example, the modulating valve 1010 is intended to initiate a return of fuel to the fuel tank 50 when the fuel having passed through one or both of the heat exchangers 1004, 1006 is at a temperature of at least 120°C (or at least 140°C; or between 120°C and 180°C; or between 140°C and 180°C) and an indication of an operating condition is provided, although this is not essential.

[0436] The operating condition may be one or more operating conditions, including, but not limited to, a proportion of sustainable aviation fuel SAF in the fuel, a thermal stability of the fuel, a coking level of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0437] The operating conditions may be detected on board, for example in flight, such as using one or more sensors and / or other measurements. Data from the one or more sensors may be provided to the fuel system electronic controller 1000 to control the operation of the modulator valve 1010. Alternatively, the operating conditions may be viewed, manually entered, or transmitted to the aircraft 1 (e.g., to the fuel system electronic engine controller 1000), either before flight or during flight.

[0438] For example, trace substances or species, naturally occurring in the fuel or added to serve as a tracer, may 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 vibration mode of a piezoelectric crystal exposed to the fuel may be used as a basis for determining various fuel characteristics, including the aromatics content of the fuel, the oxygen content of the fuel, the thermal stability of the fuel, and the coking level of the fuel, for example, by measuring the accumulation of surface deposits on the piezoelectric crystal, which will cause a change in vibration mode. Other fuel characteristics, such as sulfur content, naphthalene content, aromatic hydrocarbon content, and hydrogen / carbon ratio, may be determined by measuring substances present in the exhaust gases emitted by the gas turbine engine 10 during operation (e.g., in flight).

[0439] In this example, the modulating valve 1010 is also intended to modulate the flow of fuel along the fuel return line 1011 based, at least in part, on an amount of fuel remaining in the fuel tank, although this is not essential.

[0440] The modulating valve 1010 is further or alternatively to modulate the flow of fuel along the fuel return line 1011 such that an equilibrium temperature of the fuel in the fuel tank 50, after the fuel is returned to the fuel tank 50, does not exceed a maximum permissible temperature in the fuel tank 50 (e.g., 65°C or 100°C). The electronic fuel system controller 1000 may determine the length of time that the modulating valve 1010 may remain open and / or the amount of fuel that may be returned to the fuel tank 50 (and optionally in turn modulate the mass flow rate of the fuel returned to the fuel tank 50 using the modulating valve 1010). An example of logic that may be used to operate the modulating valve 1010 in this manner is provided below:

[0441] [Math.l] ( Mtank + MFrtt ) ' G' ( = Mtank ■ Cp ( T tank ) ■ Ttank + MFRTT • Cp ( TFRTT ) ■ Tfrtt

[0442] where Mtank is a mass of fuel in the fuel tank 50, Mfrtt is a mass of fuel returned to the fuel tank 50 along the fuel return line 1011, Te9 is an equilibrium temperature of the mixed fuel in the fuel tank 50 after the fuel is returned to the fuel tank 50, Ttat)k is a current temperature of the fuel in the fuel tank 50 before mixing, Tfrtt is the temperature of the fuel returned to the fuel tank 50 along the fuel return line 1011, and Cp is a specific heat capacity of the fuel at the respective temperature (e.g. example, Teq, Ttank, TFRTT). A mass of fuel to be returned to the fuel tank 50 may be determined based on a desired equilibrium temperature and the current temperatures of the fuel in the fuel tank 50 and the fuel returned to the fuel tank 50 along the fuel return line 1011, ensuring that the equilibrium temperature does not exceed a predetermined threshold temperature. Once the mass of fuel to be returned to the fuel tank 50 is determined, the modulating valve 1010 may be actuated accordingly to return the required mass of fuel to the fuel tank 50.

[0443] A temperature of the fuel having passed through one or both of the heat exchangers 1004, 1006 may be used to determine or be representative of a temperature of the fuel returned to the fuel tank 50. In this example, the temperature of the fuel returned to the fuel tank 50 is measured using a temperature sensor in the fuel return line 1011, although it is appreciated that the temperature of the fuel returned to the fuel tank 50 may be measured using a temperature sensor located at substantially any position downstream of one or both of the heat exchangers 1004, 1006. The temperature of the fuel in the fuel tank 50 may be measured using a temperature sensor located in the fuel tank 50 or a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006.The electronic fuel system controller 1000 is configured to control operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0444] It will be appreciated that any suitable alternative logic for controlling the operation of the modulating valve 1010 (e.g., based at least in part on an amount of fuel remaining in the fuel tank 50) may be used.

[0445] [Fig. 10] illustrates a method 2200 of operating a gas turbine engine 10. The method 2200 includes initiating 2201 a return of fuel to the fuel tank using the modulating valve 1010 when the fuel having passed through the one or more heat exchangers is at a temperature of at least 120°C, or within any of the other ranges described above with respect to the fuel system 1000 or described elsewhere herein.

[0446] In another example, the modulating valve 1010 is further or alternatively adapted to modulate the flow of fuel along the fuel return line 1011 such that a ratio between the mass of fuel returned to the tank 50 and the mass of fuel delivered to the combustion chamber 16 is between 0 and 9 under cruise conditions. The electronic fuel system controller 1000 is configured to control the operation of the modulator valve 1010. The modulator valve 1010 may be configured to modulate the flow of fuel along the fuel return line 1011 such that the ratio of the mass of fuel returned to the tank 50 to the mass of fuel delivered to the combustion chamber 16 is greater than 4 and less than or equal to 9 under cruise conditions, and preferably is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0447] SAF typically has a higher calorific value than conventional fuels such as kerosene or kerosene-based fuels. For example, SAF may have a calorific value between 43.5 MJ / kg and 44 MJ / kg, compared to a typical calorific value of 43.2 MJ / kg for a kerosene-based fuel such as Jet-A. A lower mass flow rate of SAF is therefore required to provide the same amount of fuel power to the combustion chamber as that provided by conventional fuels.

[0448] In addition, SAF typically exhibits higher thermal stability than conventional fuels and can therefore be used at a higher temperature before the fuel degrades (due to thermal degradation). SAF is therefore able to absorb more heat than conventional fuels before undergoing fuel degradation.

[0449] Returning fuel to the fuel tank 50 is typically performed when additional fuel is used for purposes other than combustion, such as to manage thermal loads in the engine 10. For example, the fuel may be used to cool the oil in the engine 10 via the heat exchangers 1004, 1006 and / or to drive one or more actuators for operating other components in the engine 10 or the aircraft 1 more generally. Because SAF can absorb more heat than conventional fuels and has a higher heating value, the inventors have determined that a different ratio of the mass of fuel returned to the fuel tank to the mass of fuel delivered to the combustion chamber may improve performance.For example, by using SAF, the increased calorific value of the fuel can result in a reduction in the mass of fuel burned, allowing a greater percentage of fuel to be returned to the tank compared to using conventional fossil fuels.

[0450] [Fig. 11] illustrates a method 2300 of operating a gas turbine engine 10. The method 2300 includes modulating 2301 the flow of fuel along the fuel return line 1011 such that a ratio of the mass of fuel returned to the fuel tank 50 to the mass of fuel delivered to the combustion chamber 16 is between 0 and 9 under cruise conditions, as described above with respect to the fuel system 1000. The ratio of the mass of fuel returned to the fuel tank 50 to the mass of fuel delivered to the combustion chamber 16 may be in or within any of the other ranges defined above or elsewhere herein.

[0451] In another example, the modulating valve 1010 is further or alternatively to prevent return of fuel having a temperature of 180°C or higher into the fuel tank 50. The electronic fuel system controller 1000 is configured to control operation of the modulating valve 1010 accordingly.

[0452] A fuel temperature is measured downstream of one or both of the heat exchangers 1004, 1006. In this example, the fuel temperature is measured using a temperature sensor in the fuel return line 1011, although it is appreciated that the fuel temperature may be measured using a temperature sensor located at substantially any position downstream of one or both of the heat exchangers 1004, 1006. A temperature of the fuel having passed through one or both of the heat exchangers 1004, 1006 may be used to determine or be representative of a fuel temperature in the fuel return line 1011. The electronic fuel system controller 1000 is configured to control operation of the modulating valve based, at least in part, on temperature data from the one or more temperature sensors located downstream of one or both of the heat exchangers 1004, 1006..

[0453] In this example, the modulating valve 1010 is also to prevent fuel from returning to the fuel tank when a temperature of the fuel in the fuel tank 50 is at a predetermined upper threshold temperature. As described above, the temperature of the fuel in the fuel tank 50 is typically maintained between -54°C (219.15 K) and 100°C (373.15 K). Preferably, it may be maintained between -54°C (219.15 K) and 65°C (338.15 K). The upper limit of the temperature of the fuel in the fuel tank 50 may be any suitable value between the pairs of temperatures of the two preceding sentences, for example, the upper limit may be 55°C (328.15 K), or 5°C (278.15 K), or 0°C (273.15 K).In some examples, the upper limit of the fuel in the fuel tank may be 0°C, 5°C, 10°C, 15°C, 20°C, 25°C, 30°C, 35°C, 40°C, 45°C, 50°C, 55°C, 60°C, 65°C, 70°C, 75°C, 80°C, 85°C, 90°C, 95°C, or 100°C.

[0454] By preventing fuel from flowing back to the fuel tank 50 when the fuel in the fuel tank 50 is at the predetermined upper threshold temperature, the threshold temperature will not be exceeded as a result of the returning fuel that has passed through one or both of the heat exchangers 1004, 1006 to the fuel tank 50. The temperature of the fuel in the fuel tank 50 may be measured using a temperature sensor located in the fuel tank 50 or a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006. The electronic fuel system controller 1000 is configured to control operation of the modulating valve 1010 based, at least in part, on temperature data from at least one of the temperature sensors providing data indicative of a temperature of the fuel returned to the fuel tank 50 and a temperature sensor providing data indicative of a temperature of the fuel in the fuel tank 50.

[0455] In this example, the modulating valve 1010 is also intended to modulate the flow of fuel along the fuel return line 1011 based, at least in part, on an amount of fuel remaining in the fuel tank, although this is not essential.

[0456] The modulating valve 1010 may be adapted to allow the return of fuel to the fuel tank 50 if a temperature of the fuel returned to the fuel tank 50 is less than 180°C and if a temperature of the fuel in the fuel tank 50 is less than the predetermined upper threshold temperature. If both of these conditions are met, the modulating valve 1010 is adapted to modulate the flow of fuel along the fuel return line 1011 such that an equilibrium temperature of the fuel in the fuel tank 50, after the return of the fuel to the fuel tank 50, does not exceed the predetermined upper threshold temperature.The electronic fuel system controller 1000 may determine the length of time that the modulator valve 1010 may remain open and / or the amount of fuel that may be returned to the fuel tank 50 (and possibly in turn modulate the mass flow rate of fuel returned to the fuel tank 50 using the modulator valve 1010). An example of logic that may be used to operate the modulator valve 1010 in this manner is provided below: .

[0457] [Math.2] (Mtank + MfRTT > 'CpÇl a / ) -1^ = Mlank 'Cp^Ttank ) ■ Tumk + MFRTT TFKTT ) • T FRTT

[0458] where Mtank is a mass of fuel in the fuel tank 50, MFRTT is a mass of fuel returned to the fuel tank 50 along the fuel return line 1011, is an equilibrium temperature of the mixed fuel in the fuel tank 50 after the fuel is returned to the fuel tank 50, Tt(mk is a current temperature of the fuel in the fuel tank 50 before mixing, Tfrtt is a current temperature of the fuel returned to the fuel tank 50 along the fuel return line 1011, and Cp is a specific heat capacity of the fuel at the respective temperature (e.g.,^, Ttmik, A mass of fuel to be returned to the fuel tank 50 may be determined based on a desired equilibrium temperature and the current temperatures of the fuel in the fuel tank 50 and the fuel returned to the fuel tank 50 along the fuel return line 1011, ensuring that the equilibrium temperature does not exceed a predetermined upper threshold temperature for the fuel in the fuel tank 50. Once the mass of fuel to be returned to the fuel tank 50 is determined, the modulating valve 1010 may be actuated accordingly to return the required mass of fuel to the fuel tank 50.

[0459] A temperature of the fuel having passed through one or both of the heat exchangers 1004, 1006 may be used to determine or be representative of a temperature of the fuel returned to the fuel tank 50. In this example, the temperature of the fuel returned to the fuel tank 50 is measured using a temperature sensor in the fuel return line 1011, although it is appreciated that the temperature of the fuel returned to the fuel tank 50 may be measured using a temperature sensor located at substantially any position downstream of one or both of the heat exchangers 1004, 1006. The temperature of the fuel in the fuel tank 50 may be measured using a temperature sensor located in the fuel tank 50 or a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006.The electronic fuel system controller 1000 is configured to control operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0460] It will be appreciated that any suitable alternative logic for controlling the operation of the modulating valve 1010 (e.g., based at least in part on an amount of fuel remaining in the fuel tank 50) may be used.

[0461] [Fig. 12] illustrates a method 2400 of operating a gas turbine engine 10. The method 2400 includes preventing 2401 a return of fuel having a temperature of 180°C or higher to the fuel tank 50 using the modulating valve 1010, as described above with respect to the fuel system 1000. The method 2400 may include preventing 2401 a return of fuel having a temperature within any of the ranges defined above or elsewhere herein to the fuel tank 50 using the modulating valve 1010. The method may include preventing fuel from returning if it would exceed the upper limit of the tank temperature defined above or as defined elsewhere herein.

[0462] The fuel system 1500 may be configured to perform one or more of the methods 2000, 2100, 2200, 2300, 2400. It may also perform the method of [Fig. 13] described below if it is equipped with a suitable fuel recirculation line.

[0463] Another example of a fuel system 1100 including a fuel flow path from the fuel tank 50 to other components of the fuel system 1100 is shown schematically in [Fig. 6]. The fuel system 1100 is substantially similar to the fuel system 1000 described above with respect to [Fig. 5], with like reference numerals indicating like elements.

[0464] The fuel system 1100 includes a fuel recirculation line 1111. The fuel recirculation line 1111 is for recirculating fuel from a first point in the fuel flow path to a second point in the fuel flow path. In this example, the fuel flow path may be referred to as the primary fuel flow path because it extends from the fuel tank to the combustion chamber. The second point in the fuel flow path is upstream of the first point. The first point corresponds to an inlet of the fuel recirculation line 1111. The second point corresponds to an outlet of the fuel recirculation line 1111. The fuel system 1100 includes a modulating valve 1110 for modulating a flow of fuel along the fuel recirculation line 1111, between the inlet and the outlet of the fuel recirculation line 1111.The fuel system 1100 includes a temperature sensor located in the fuel recirculation line 1111, although this is not essential. The fuel system 1100 may also include a fuel return line (not shown for clarity) as described above with respect to the fuel system 1000 shown in [Fig. 5], although this is not essential.

[0465] In the illustrated example, the first point or inlet of the fuel recirculation line 1111 is located downstream of the primary fuel-oil heat exchanger 1106 in the fuel flow path. The first point is located downstream of the fuel pump 1103. The modulator valve 1110 is also located in the fuel flow path, although it will be appreciated that the modulator valve 1110 may instead be located along the fuel recirculation line 1111. The second point or outlet of the fuel recirculation line 1111 is located upstream of the primary fuel-oil heat exchanger 1106.

[0466] The inlet and outlet of the fuel recirculation line 1111 may alternatively be located at any suitable position in the fuel flow path. For example, the inlet of the fuel recirculation line 1111 may be located upstream of the fuel pump 1003, or downstream of the secondary fuel-oil heat exchanger 1104, or downstream of the fuel pump 1102. The outlet of the fuel recirculation line 1111 may be located upstream of the fuel pump 1103, or upstream of the secondary fuel-oil heat exchanger 1104, or upstream of the fuel pump 1102. When the outlet of the fuel recirculation line 1111 is located upstream of the fuel pump 1102, the inlet of the recirculation line may be upstream of the pump 1103.It will also be appreciated that the fuel system 1100 may alternatively comprise a single fuel-oil heat exchanger or may comprise more than two fuel-oil heat exchangers, and the first point or second point of the fuel recirculation line 1111 may be located at any suitable position (upstream or downstream, directly or indirectly with one or more components intervening in the main fuel flow path) respective to the one or more heat exchangers. A fuel recirculation line 1111 may be provided for any given fuel-oil heat exchanger.

[0467] The fuel recirculation line 1111 is to allow a portion of the fuel to flow along at least a portion of the fuel flow path a plurality of times (i.e., to recirculate a portion of the fuel along at least a portion of the fuel flow path). In the illustrated example, the modulating valve 1110 is to modulate the flow of fuel along the fuel recirculation line 1111 such that a quantity of fuel passes through the primary fuel-oil heat exchanger 1106 a plurality of times before being delivered to the combustion chamber 16. This may allow additional heat to be transferred into the same fuel and increase the temperature of the fuel before it is delivered to the combustion chamber 16.A fuel temperature may be increased to at least 120°C before fuel is delivered to the combustion chamber 16. Recirculating the fuel through the fuel recirculation line 1111 using the modulator valve 1110 may therefore allow hotter fuel to be delivered to the combustion chamber 16, which may increase the efficiency of the engine 10. Further, an amount (e.g., mass) of fuel used to manage thermal loads in the engine 10 (e.g., via the heat exchangers 1104, 1106) may be reduced, which in turn may reduce (or eliminate, depending on operating conditions) an amount of fuel that needs to be returned. to the fuel tank 50 when managing thermal loads in the engine 10. Reducing or eliminating an amount of fuel returned to the fuel tank 50 may also allow the fuel in the fuel tank 50 to be cooler before entering the fuel flow path, which may improve cooling of the oil in the fuel-oil heat exchangers 1104, 1106.

[0468] The modulating valve 1110 and / or the inlet of the fuel recirculation pipe 1111 may be located downstream of the fuel pump 1103, such that the recirculated fuel passes through both the primary fuel-oil heat exchanger 1106 and through the fuel pump 1103 multiple times rather than passing only through the primary heat exchanger 1106. The recirculated fuel through the fuel pump 1103 may allow for more adjustable control of the fuel flow to the combustion chamber 16 for a given shaft speed of the engine 10, noting that the pump speed (or a limited group of pump speed options) is often controlled by the shaft speed (e.g., the pump speed may be proportionally related to the shaft speed).For example, if a fuel flow rate required to maintain current aircraft thrust is less than the pump speed for the required shaft speed, a greater proportion of fuel may be recirculated through the pump such that the flow rate through the pump is higher than the flow rate to the combustion chamber 16. A lower fuel flow rate to the combustion chamber 16 may also be provided for a higher heating value fuel, without changing the pump speed, but rather by increasing the amount of recirculation. Typically, there is no wasted fuel, as it is always recirculated and burned in the combustion chamber.

[0469] In the illustrated example, the modulating valve 1110 is intended to actively control an amount of fuel recirculated along the fuel recirculation line 1111. The electronic fuel system controller 1000 is configured to control the operation of the modulating valve 1110. The active control may be performed based on one or more parameters such as: • heart shaft speed and engine thrust demand; • current fuel temperature at one or more locations, for example along the fuel flow path (such as downstream of one or more heat exchangers) and / or along the fuel recirculation line; • calorific value of fuel; • fan rotation speed; • fuel flow to the combustion chamber (commonly referred to as WFE - weight of main engine fuel flow); • fuel viscosity; and • main or engine fuel pump speed, or speed options.

[0470] Additionally or alternatively, a speed of the fuel pump 1103 may be adjusted, either by accelerating the fuel flow rate (and thereby decreasing the heat transfer per unit volume of fuel passing through the heat exchangers 1104, 1106), or by reducing the fuel flow rate (and thereby increasing the heat transfer per unit volume of fuel passing through the heat exchangers 1104, 1106).

[0471] Alternatively, control of the amount of fuel recirculated along the fuel recirculation line 1111 using the modulating valve 1110 may not be active. Rather, a set or fixed proportion of fuel in the main fuel flow path may be recirculated along the fuel recirculation line 1111 via the modulating valve 1110, for example, the ratio of the mass of fuel recirculated to the mass of fuel delivered to the combustion chamber may be between 0 and 9. More preferably, the ratio of the mass of fuel recirculated to the mass of fuel delivered to the combustion chamber may be greater than 4 and less than or equal to 9 under cruise conditions, and even more preferably greater than 4 and less than or equal to 4.9 under cruise conditions.It may therefore be the same as the ratio of fuel returned to the tank as defined above, such that everything described herein that applies to fuel returned to the tank may also apply to the present example in which the fuel is instead recirculated along the fuel recirculation line 1111.

[0472] In some examples, the fuel recirculation line 1111 is intended to supply fuel to one or more additional aircraft and / or engine mechanisms (not shown) before joining the fuel flow path. The one or more additional aircraft mechanisms may therefore be located on the fuel recirculation line 1111. The additional aircraft mechanisms may include a nacelle anti-icing system, actuators, a turbine case cooling system, or any other suitable aircraft mechanism.

[0473] [Fig. 13] illustrates a method 2500 of operating a gas turbine engine 10. The method 2500 includes modulating 2501 the flow of fuel along the fuel recirculation line 1111, as described above with respect to the fuel system 1100.

[0474] Another example of a fuel system 1200 including a fuel flow path from the fuel tank 50 to other components of the fuel system 1200 is shown schematically in [Fig. 7]. In the example shown, the engine 10 is a direct drive engine, as shown in [Fig. 4]. The fuel system 1200 of [Fig. 7], however, may be used with other engine architectures, such as a geared architecture. The fuel system 1200 includes both the fuel supply system (including the fuel tank 50 and the pump 1202) that supplies fuel to the engine 10 and the fuel management system 1600 (which operates within the engine 10) of the aircraft 1. In the described implementation, each engine 10 has its own fuel management system 1600.In other implementations, a single fuel management system 1600 may manage the fuel supply to multiple engines, and may, for example, include a duplicate of various of the elements shown in [Fig. 6] for the other engine(s).

[0475] Fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low pressure fuel feed pump 1202. The fuel then flows through a primary fuel-to-oil heat exchanger 1204 before flowing through an engine fuel pump 1203. The engine fuel pump 1203 can be described as a primary or high pressure fuel pump. At least a portion of the fuel then passes through a secondary fuel-oil heat exchanger 1206 via a line 1207 that branches off the primary fuel flow path (between the fuel tank 50 and the combustion chamber 16), and at least a portion of the fuel flows to the combustion chamber 16 without passing through the secondary fuel-oil heat exchanger 1206. The primary fuel-oil heat exchanger 1204 may be referred to as the primary fuel-oil heat exchanger.The secondary fuel-oil heat exchanger 1206 may be referred to as a slave fuel-oil heat exchanger. The fuel management system 1600 is arranged such that fuel reaches the primary fuel-oil heat exchanger 1204 before the secondary fuel-oil heat exchanger 1206.

[0476] In the example shown in [Fig.7], from the secondary fuel-oil heat exchanger 1206, the portion of fuel that has flowed therethrough then flows to the combustion chamber 16 of the gas turbine engine 10, joining the fuel that has flowed through the primary fuel-oil heat exchanger 1204 only.

[0477] The primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 are designed such that an oil flow is also carried therethrough. The fuel-oil heat exchanger The primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 are designed such that heat can be transferred between the oil and the fuel flowing therethrough. Under cruise conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 respectively. In this manner, the primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 are designed to transfer thermal energy from an oil flow to a fuel flow flowing therethrough under cruise conditions.The fuel system 1200 includes an electronic control device configured to control the operation of the heat exchangers 1204, 1206.

[0478] In various arrangements of the fuel management system 1600, the oil flows through the secondary fuel-oil heat exchanger 1206 before passing through the primary fuel-oil heat exchanger 1204, and does not flow through engine components that would increase its temperature therebetween. The oil is therefore hotter at the inlet of the secondary fuel-oil heat exchanger 1206 than at the inlet of the primary fuel-oil heat exchanger 1204. In contrast, the fuel flows through the primary fuel-oil heat exchanger 1204 before flowing through the secondary fuel-oil heat exchanger 1206. In this arrangement, the fuel temperature exiting the secondary fuel-oil heat exchanger 1206 is higher than the fuel temperature exiting the primary fuel-oil heat exchanger 1204.The fuel pump 1203 may alternatively be placed upstream of the secondary fuel-oil heat exchanger 1206 (or upstream of where the line 1207 joins the main fuel flow path), which means that it is not exposed to these even higher fuel temperatures.

[0479] The fuel management system 1600 may be provided with a bypass to allow a portion of the fuel to bypass passing through the respective heat exchanger, for example in the form of a bypass pipe with an inlet located upstream of the primary fuel-oil heat exchanger 1204 and an outlet located downstream of the primary fuel-oil heat exchanger 1204 and upstream of the secondary fuel-oil heat exchanger 1206, as described above with respect to [Fig. 5]. A valve may determine what proportion of the fuel passes through the heat exchanger 1204 and what proportion passes through the bypass pipe.

[0480] The fuel system 1200 further includes a modulator valve 1210 located downstream of the primary fuel-oil heat exchanger 1204 and configured to divert or direct at least a portion of the fuel that has left the primary fuel-oil heat exchanger 1204 back to the fuel tank 50 via a fuel return line 1211. The modulator valve 1210 is configured to modulate a flow of fuel along the fuel return line 1211. The modulator valve 1210 can determine what proportion of the fuel is returned to the fuel tank 50, and what proportion continues to the combustion chamber 16. It will be appreciated that the modulator valve 1210 can be located at substantially any position downstream of at least one of the heat exchangers 1204, 1206.For example, the modulator valve 1210 may alternatively be located upstream of the secondary heat exchanger 1206 (e.g., upstream of where the line 1207 joins the primary fuel flow path). It will also be appreciated that the fuel system 1200 may alternatively include a single fuel-oil heat exchanger or may include more than two fuel-oil heat exchangers, and that the modulator valve 1210 may be located downstream of at least one of the heat exchangers (either directly downstream or with one or more intermediate components of the fuel system 1200 between the modulator valve 1210 and the heat exchanger). The primary and secondary heat exchangers described above are only exemplary and may provide cooling for any engine systems or components.

[0481] In the illustrated example, the modulating valve 1210 is located downstream of the primary fuel-oil heat exchanger 1204, the primary fuel pump 1203, the inlet of the line 1207 leading to the secondary fuel-oil heat exchanger 1206, and the outlet of the line 1207 joining the primary fuel flow path from the secondary fuel-oil heat exchanger 1206. It will be appreciated that the modulating valve 1210 may alternatively be located upstream of the fuel pump 1203.

[0482] The fuel system 1200 includes a plurality of temperature sensors (indicated by the circular symbols surrounding the capital letter T) configured to measure a temperature of the fuel at different locations in the fuel system 1200. In the illustrated example, the fuel system 1200 includes a temperature sensor located in the fuel tank 50 for measuring a temperature of the fuel in the fuel tank 50. Additionally or alternatively, the fuel system 1200 may include a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers. heat 1204, 1206, which may be used to determine or be representative of a fuel temperature in the fuel tank 50.

[0483] In the illustrated example, the fuel system 1200 includes two temperature sensors located downstream of the heat exchangers 1204, 1206. One of the temperature sensors is located in the fuel return line 1211, while the other temperature sensor is located downstream of the modulating valve 1210 and upstream of the combustion chamber 16. Either or both of the temperature sensors may be used to determine or represent a temperature of the fuel delivered to the combustion chamber 16 and / or a temperature of the fuel returned to the fuel tank 50. It will be appreciated that only one of the temperature sensors located downstream of the heat exchangers 1204, 1206 may be provided.Depending on the position of the modulating valve 1210 relative to one or more heat exchangers (e.g., the primary heat exchanger 1204 and / or the secondary heat exchanger 1206), it will be appreciated that the temperature sensor(s) may be placed at any suitable location downstream of the heat exchanger(s) in order to measure temperature data relating to (or provide temperature data representative of) a temperature of the fuel delivered to the combustion chamber 16 and / or a temperature of the fuel returned to the fuel tank 50.

[0484] Returning fuel to the fuel tank 50 provides a mechanism for controlling fuel flow of the fuel system 1200, for example to manage thermal loads in the engine 10 and / or to control fuel temperature at different locations in the system 1200 (such as in the fuel tank 50 or at the inlet of the combustion chamber 16).

[0485] The modulating valve 1210 may be adapted to modulate the flow of fuel along the fuel return line 1211 to the fuel tank 50 in any suitable manner, for example as described above with respect to the fuel system 1000 illustrated in [Fig. 5]. The fuel system 1200 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulating valve 1210. It will also be appreciated that any one or more of the methods 2000, 2100, 2200, 2300, 2400 illustrated in FIGS. 8-12 and described above with respect to the fuel system 1000 may also be performed using the fuel system 1200 illustrated in [Fig. 7]. The fuel system 1200 may also perform the method 2500 of [Fig. 13] if equipped with a suitable fuel recirculation line, as described below.

[0486] Another example fuel system 1300 including a fuel flow path from the fuel tank 50 to other components of the fuel system fuel system 1300 is schematically shown in [Fig. 14]. The fuel system 1300 is substantially similar to the fuel system 1200 described above with respect to [Fig. 7], with similar reference numerals indicating similar elements.

[0487] The fuel system 1300 includes a fuel recirculation line 1311. The fuel recirculation line 1311 is for recirculating fuel from a first point in the fuel flow path to a second point in the fuel flow path. In this example, the fuel flow path is a main fuel flow path that extends from the fuel tank to the combustion chamber. The second point in the fuel flow path is upstream of the first point. The first point corresponds to an inlet of the fuel recirculation line 1311. The second point corresponds to an outlet of the fuel recirculation line. The fuel system 1300 includes a modulating valve 1310 for modulating a flow of fuel along the fuel recirculation line 1311, between the inlet and the outlet of the fuel recirculation line 1311.The fuel system 1300 includes a temperature sensor located in the fuel recirculation line 1311, although this is not essential. The fuel system 1300 may also include a fuel return line (not shown for clarity) as described above with respect to the fuel system 1200 shown in [Fig. 7], although this is not essential.

[0488] In the illustrated example, the first point or inlet of the fuel recirculation line 1311 is located downstream of the primary fuel-oil heat exchanger 1304 on the fuel flow path. The first point is located downstream of the fuel pump 1303. The first point is located downstream of the inlet of the line 1307 leading to the secondary fuel-oil heat exchanger 1306, and downstream of the outlet of the line 1307 joining the primary fuel flow path from the secondary fuel-oil heat exchanger 1306 (between the fuel tank 50 and the combustion chamber 16). The first point is therefore downstream of the flow paths to and from the secondary heat exchanger 1306. The second point is located upstream of the flow paths to and from the secondary heat exchanger.The modulator valve 1310 is also located on the primary fuel flow path, although it will be appreciated that the modulator valve 1310 may instead be located at another position on the fuel flow path or along the fuel recirculation line 1311. The second point or outlet of the fuel recirculation line 1311 is located upstream of the primary fuel-oil heat exchanger 1304.

[0489] The inlet and outlet of the fuel recirculation line 1311 may alternatively be located at any suitable position in the fuel flow path. For example, the inlet of the fuel recirculation line 1311 may be located downstream of the primary heat exchanger 1304 but upstream of the fuel pump 1303, or downstream of the primary heat exchanger 1304 and the fuel pump 1303 but upstream of the inlet of the line 1307 leading to the secondary heat exchanger 1306, or downstream of the outlet of the line 1307 joining the primary fuel flow path from the secondary heat exchanger 1306. The outlet of the fuel recirculation line 1311 may be located upstream of the inlet of the line 1307 leading to the secondary heat exchanger 1306, or upstream of the fuel pump 1303, or upstream of the fuel pump 1302.It will also be appreciated that the fuel system 1300 may alternatively comprise a single fuel-oil heat exchanger or may comprise more than two fuel-oil heat exchangers, and the first point or second point of the fuel recirculation line 1311 may be located at any suitable position (upstream or downstream, directly or indirectly with one or more components intervening in the fuel flow path) respective to the one or more heat exchangers. A fuel recirculation line 1311 may be provided for any given fuel-oil heat exchanger.

[0490] The fuel recirculation line 1311 is to allow a portion of the fuel to flow along at least a portion of the fuel flow path a plurality of times (i.e., to recirculate a portion of the fuel along at least a portion of the fuel flow path) before being delivered to the combustion chamber 16. In the illustrated example, the modulating valve 1310 is to modulate the flow of fuel along the fuel recirculation line 1311 such that an amount of fuel passes through the primary fuel-oil heat exchanger 1304 a plurality of times before being delivered to the combustion chamber 16. This may allow additional heat to be transferred into the same fuel and increase the temperature of the fuel before it is delivered to the combustion chamber 16.A fuel temperature may be increased to at least 120°C before fuel is delivered to the combustion chamber 16. In some examples, the fuel temperature may be increased to at least 140°C before fuel is delivered to the combustion chamber 16, or within the range between 120°C and 180°C or between 140°C and 180°C. Recirculating the fuel through the fuel recirculation line 1311 using the modulating valve 1310 may therefore allow hotter fuel to be delivered to the combustion chamber. 16, which may increase the efficiency of the engine 10. Further, an amount (e.g., mass) of fuel used to manage thermal loads in the engine 10 (e.g., via the heat exchangers 1304, 1306) may be reduced, which in turn may reduce (or eliminate, depending on operating conditions) an amount of fuel that must be returned to the fuel tank 50 when managing thermal loads in the engine 10. Reducing or eliminating an amount of fuel returned to the fuel tank 50 may also allow the fuel in the fuel tank 50 to be cooler before entering the fuel flow path, which may improve cooling of the oil in the fuel-oil heat exchangers 1304, 1306.

[0491] In the illustrated example, the modulating valve 1310 and / or the inlet of the fuel recirculation pipe 1311 is also located downstream of the fuel pump 1303, such that the recirculated fuel passes through both the primary fuel-oil heat exchanger 1304 and the fuel pump 1303 multiple times. The recirculated fuel through the fuel pump 1303 may allow for more adjustable control of the fuel flow to the combustion chamber 16 for a given shaft speed of the engine 10, noting that the pump speed (or a limited group of pump speed options) is often controlled by the shaft speed (e.g., the pump speed may be proportionally related to the shaft speed).For example, if a fuel flow rate required to maintain current aircraft thrust is less than the pump speed for the required shaft speed, a greater proportion of fuel may be recirculated through the pump such that the flow rate through the pump is higher than the flow rate to the combustion chamber 16. A lower fuel flow rate to the combustion chamber 16 may also be provided for a higher heating value fuel, without changing the pump speed, but rather by increasing the amount of recirculation.

[0492] The modulating valve 1310 may be adapted to modulate the flow of fuel along the fuel recirculation line 1311 in any suitable manner, for example as described above with respect to the fuel system 1200 illustrated in [Fig. 7]. The fuel system 1300 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulating valve 1310. It will also be appreciated that the method 2500 illustrated in [Fig. 13] and described above with respect to the fuel system 1100 may also be performed using the fuel system 1300 illustrated in [Fig. 14].

[0493] In some examples, the fuel recirculation line 1311 is for supplying fuel to one or more aircraft and / or engine mechanisms. additional (not shown) before joining the fuel flow path. The additional aircraft mechanism(s) may therefore be located on the fuel recirculation line 1311. The additional aircraft mechanisms may include a nacelle anti-icing system, actuators, a turbine case cooling system, or any other suitable aircraft mechanism.

[0494] Another example of fuel systems 1400, 1700 including a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 are shown schematically in Figures 15 and 16. The fuel system 1400 is similar to the fuel systems 1000, 1100 described above with respect to Figures 5 and 6, with like reference numerals indicating like elements. The fuel system 1700 is similar to the fuel systems 1200, 1300 described above with respect to Figures 7 and 14, with like reference numerals indicating like elements.

[0495] In fuel systems 1400, 1700, a modulating valve 1410, 1710 is provided to divert or direct at least a portion of the fuel that has exited a fuel-oil heat exchanger back to the fuel tank 50 and / or to recirculate the fuel to an upstream location in the fuel flow path. It may therefore perform the method described in one or more of FIGS. 8-13.

[0496] In the example of [Fig. 15], the modulator valve 1410 of the fuel system 1400 is located downstream of the secondary fuel-oil heat exchanger 1406 and the pump 1403. The modulator valve 1410 is intended to divert or direct at least a portion of the fuel from the pump 1403 back to the fuel tank 50 via a fuel return line 1411b, and is also intended to recirculate the fuel from the pump 1403 to a point upstream of the secondary heat exchanger 1406 via a fuel recirculation line 1411a. It will also be appreciated that the modulator valve 1410 may alternatively be placed at any suitable location in the fuel system 1400, for example in the fuel flow path.

[0497] The modulating valve 1410 may be adapted to modulate the flow of fuel along the fuel recirculation line 1411a and the fuel return line 1411b in any suitable manner, for example as described above with respect to the fuel systems 1000, 1100, 1200, 1300 illustrated in Figures 5, 6, 7 and 14. The fuel system 1400 includes an electronic controller configured to control the operation (e.g., the opening and closing) of the modulating valve 1410. It will also be borne in mind that the methods 2000, 2100, 2200, 2300, 2400, 2500 illustrated in FIGS. 8 to 13 and described above with respect to the fuel systems 1000, 1100, 1200, 1300 can also be carried out using the fuel system 1400 illustrated in [Fig.15].

[0498] In the example of [Fig.16], the modulating valve 1710 of the fuel system 1700 is located downstream of the secondary fuel-oil heat exchanger 1706. The fuel exiting the secondary heat exchanger 1706 does not join the main fuel flow path (between the fuel tank 50 and the combustion chamber 16). The modulating valve 1710 is to divert or direct at least a portion of the fuel exiting the secondary heat exchanger 1706 back to the fuel tank 50 via a fuel return line 1711b, and is also to recirculate the fuel exiting the secondary heat exchanger 1706 to a point upstream of the primary fuel-oil heat exchanger 1704 in the primary fuel flow path via a fuel recirculation line 1711a.It will also be appreciated that the modulating valve 1710 may alternatively be placed at any suitable location, for example on the main fuel flow path, such that fuel exiting the secondary heat exchanger 1706 may join the main fuel flow path before encountering the modulating valve 1710.

[0499] The modulating valve 1710 may be adapted to modulate the flow of fuel along the fuel recirculation line 1711a and the fuel return line 1711b in any suitable manner, for example as described above with respect to the fuel systems 1000, 1100, 1200, 1300 illustrated in Figures 5, 6, 7 and 14. The fuel system 1700 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulating valve 1710. It will also be appreciated that the methods 2000, 2100, 2200, 2300, 2400, 2500 illustrated in Figures 8-13 and described above with respect to the fuel systems 1000, 1100, 1200, 1300 can also be implemented using the fuel system 1700 shown in [Fig. 16],

[0500] The temperatures of fuel delivered to the combustion chamber under cruise conditions, or at which fuel return is initiated, defined at any point herein, may be defined as an average over at least 5 minutes, 10 minutes, or 30 minutes, under steady-state cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in fuel temperature during operation, often a rise in temperature. Each fluctuation may not last more than 5 minutes.

[0501] [Fig. 17] shows an aircraft 1 on which two gas turbine engines 10 of the present disclosure are mounted, one on each wing. Fuel F is supplied from fuel tank 50 to the gas turbine engines 10. In this example, fuel tank 50 comprises a set of interconnected fuel tanks. In the example of [Fig. 17], the fuel tank consists of a main fuel tank located in the fuselage of the aircraft and a smaller fuel tank located in each wing. In other examples, an aircraft 1 may have only a single fuel tank - it will be appreciated that many different tank arrangements are possible and that the illustrated examples are provided for ease of description and are not intended to be limiting.

[0502] It will be understood that the invention is not limited to the examples described above and that various modifications and improvements may be made without departing from the concepts described herein. Except where mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Claims

1. A gas turbine engine (10) for an aircraft (1), comprising: a combustion chamber (16); a fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) arranged to receive fuel from a fuel tank (50) on board the aircraft (1) and transfer heat from the oil to the fuel; a fuel return line (1011; 1211; 1411b; 1711b) arranged to return at least a portion of the fuel that has passed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulating valve (1010; 1210; 1410; 1710) arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b;1711b) such that a ratio between AT at cruising speed and AT at startup is less than 0.56, wherein AT is a ratio between a temperature, in kelvin, of the fuel in the fuel tank (50) and a temperature, in kelvin, of the fuel delivered to the combustion chamber (16).;

2. A gas turbine engine (10) according to claim 1, wherein the modulating valve (1010; 1210; 1410; 1710) is arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) such that the ratio of AT at cruise to AT at startup is less than 0.

53.

3. A gas turbine engine (10) according to claim 1, wherein the modulating valve (1010; 1210; 1410; 1710) is arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) such that the ratio between AT at cruise and AT at start is between 0.48 and 0.56, or optionally between 0.48 and 0.

53.

4. A gas turbine engine (10) for an aircraft (1), comprising: a combustion chamber (16); a fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) arranged to receive fuel from a fuel tank (50) on board the aircraft (1) and transfer heat from the oil to the fuel; a fuel return line (1011; 1211; 1411b; 1711b) arranged to return at least a portion of the fuel that has passed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulating valve (1010; 1210; 1410; 1710) arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) such that a ratio between AT at cruise and AT at startup is between 0.48 and 1.88, wherein AT is a ratio between a temperature, in kelvin, of the fuel in the fuel tank (50) and a temperature, in kelvin, of the fuel delivered to the combustion chamber (16).

5. A gas turbine engine (10) according to claim 4, wherein the modulating valve (1010; 1210; 1410; 1710) is arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) such that the ratio between AT at cruise and AT at start is: a) between 0.48 and 1.71; b) between 0.48 and 1.70 and optionally if AT at start is determined for starting the aircraft when it is on the ground; c) between 0.48 and 1.55 and optionally if AT at start is determined for starting the aircraft when it is on the ground; d) between 0.48 and 1.88 and optionally if AT at start is determined for starting the aircraft when it is in flight; or (e) between 0.48 and 1.71 and optionally if AT at start is determined for the start of the aircraft when it is in flight.

6. A gas turbine engine (10) according to any preceding claim, wherein the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) is arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature to at least 120°C, and preferably to at least 140°C, at the inlet of the combustion chamber (16), and optionally so as to raise the fuel temperature to between 120°C and 180°C, and preferably between 140°C and 180°C, at the inlet of the combustion chamber (16).

7. A gas turbine engine (10) according to any preceding claim, wherein the gas turbine engine (10) further comprises: i) a first temperature sensor located downstream of the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706); and ii) and is configured to receive information from a second temperature sensor located in the fuel tank (50); and optionally: a) wherein the first temperature sensor is located in the fuel return line (1011; 1211; 1411b; 1711b); and / or b) wherein the modulating valve (1010; 1210; 1410; 1710) is arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) based on temperature data from the first temperature sensor and the second temperature sensor.

8. A gas turbine engine (10) according to any preceding claim, wherein the fuel temperature in the fuel tank (50) is between -54°C and 100°C, and preferably is between -54°C and 65°C, between -54°C and 55°C or between -54°C and 5°C.

9. A method (2100) of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber (16); a fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) arranged to receive fuel from a fuel tank (50) on board the aircraft (1) and transfer heat from the oil to the fuel; a fuel return line (1011; 1211; 1411b; 1711b) arranged to return at least a portion of the fuel that has passed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulating valve (1010; 1210; 1410; 1710) arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b); wherein the method comprises modulating (2101) the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) using the modulating valve (1010; 1210; 1410; 1710) such that a ratio of AT at cruise to AT at startup is less than 0.56, wherein AT is a ratio of a temperature, in kelvin, fuel in the fuel tank (50) and a temperature, in kelvin, of the fuel delivered to the combustion chamber (16).

10. The method (2100) of claim 9, wherein modulating (2101) the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulating valve (1010; 1210; 1410; 1710) comprises modulating the fuel flow such that the ratio of AT at cruise to AT at startup is less than 0.53, or optionally between 0.48 and 0.56 or between 0.48 and 0.

53.

11. A method (2100) of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber (16); a fuel-oil heat exchanger (1004; 1006; 1104; 1106; 1204; 1206; 1304; 1306; 1404; 1406; 1704; 1706) arranged to receive fuel from a fuel tank (50) on board the aircraft (1) and transfer heat from the oil to the fuel; a fuel return line (1011; 1211; 1411b; 1711b) arranged to return at least a portion of the fuel that has passed through the heat exchanger (1004; 1006; 1104; 1106; 1204; 1206; 1304; 1306; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulating valve (1010; 1210; 1410; 1710) for modulating the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b);wherein the method comprises modulating (2101) the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) using the modulating valve (1010; 1210; 1410; 1710) such that a ratio between AT at cruise and AT at startup is between 0.48 and 1.88, wherein AT is a ratio between a temperature, in kelvin, of the fuel in the fuel tank (50) and a temperature, in kelvin, of the fuel delivered to the combustion chamber (16).;

12. The method (2100) of claim 11, wherein modulating (2101) the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulating valve (1010; 1210; 1410; 1710) comprises modulating the fuel flow such that the ratio of AT at cruise to AT at startup is: a) between 0.48 and 1.71; b) between 0.48 and 1.70 and optionally if AT at start is determined for starting the aircraft while it is on the ground; c) between 0.48 and 1.55 and optionally if AT at start is determined for starting the aircraft while it is on the ground; d) between 0.48 and 1.88 and optionally if AT at start is determined for starting the aircraft while it is in flight; or e) between 0.48 and 1.71 and optionally if AT at start is determined for starting the aircraft while it is in flight.

13. A method (2100) according to any one of claims 9 to 12, comprising transferring heat from the oil to the fuel using the fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) so as to raise the fuel temperature to at least 120°C, and preferably at least 140°C, at the inlet of the combustion chamber (16), and optionally raising the fuel temperature to between 120°C and 180°C, and preferably between 140°C and 180°C, at the inlet of the combustion chamber (16).

14. The method (2100) of any one of claims 9 to 13, wherein the gas turbine engine (10) further comprises: i) a first temperature sensor located downstream of the heat exchanger; and ii) is configured to receive information from a second temperature sensor located in the fuel tank; and optionally: a) wherein the first temperature sensor is located in the fuel return line (1011; 1211; 1411b; 1711b); and / or b) the modulating valve (1010; 1210; 1410; 1710) is arranged to modulate the flow of fuel along the fuel return line (1011; 1211; 1411b; 1711b) based on temperature data from the first temperature sensor and the second temperature sensor.

15. A method (2100) according to any one of claims 9 to 14, wherein the temperature of the fuel in the fuel tank (50) is between -54°C and 100°C, and preferably between -54°C and 65°C, between -54°C and 55°C or between -54°C and 5°C.

Images