GAS TURBINE OPERATION

Abstract

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 flowed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulator valve (1010; 1210; 1410; 1710) arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) such that a ratio of the fuel mass returned to the tank 50 to the fuel mass delivered to the combustion chamber 16 is between 0 and 9. A method (2300) for operating a gas turbine engine is also disclosed.

Description

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

[0002] The aviation industry is expected to evolve toward the use of fuels different from the conventional kerosene-based jet fuels currently in widespread use. It is desirable to manage heat and fuel within the engine according to the type of fuel used. This may allow for the advantageous use of fuel properties not available in kerosene-based jet fuel.

[0003] According to a first aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow 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.

[0004] The modulator valve may be arranged to modulate the fuel flow 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.

[0005] The modulator valve may be arranged to modulate the fuel flow 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.

[0006] The modulator valve may be arranged to modulate the fuel flow 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.

[0007] According to a second aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line; wherein the method comprises modulating the fuel flow along the fuel return line using the modulator 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.

[0008] The method may include modulating the fuel flow along the fuel return line using the modulator 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.

[0009] The method may include modulating the fuel flow along the fuel return line using the modulator 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.

[0010] The method may include modulating the fuel flow along the fuel return line using the modulator 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.

[0011] According to a third aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line such that the ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 1.00.

[0012] The modulator valve may be arranged to modulate the fuel flow 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: a) is between 0.48 and 0.90; b) is between 0.48 and 0.82; c) is between 0.56 and 1.00; d) is between 0.56 and 0.82; or e) is between 0.56 and 0.75.

[0013] According to a fourth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line; wherein the method comprises modulating the fuel flow along the fuel return line using the modulator 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.

[0014] The method may include modulating the fuel flow along the fuel return line using the modulator 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: a) is between 0.48 and 0.90; b) is between 0.48 and 0.82; c) is between 0.56 and 1.00; d) is between 0.56 and 0.82; or e) is between 0.56 and 0.75.

[0015] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line such that a ratio of ΔT at cruise to ΔT at takeoff is less than 0.56, where ΔT is a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin.

[0016] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT at cruise to ΔT at takeoff is less than 0.53.

[0017] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 0.56.

[0018] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 0.53.

[0019] According to a sixth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line such that a ratio of ΔT at cruise to ΔT at takeoff is less than 0.56, where ΔT is a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin.

[0020] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of ΔT at cruise to ΔT at takeoff is less than 0.53.

[0021] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 0.56.

[0022] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 0.53.

[0023] According to a seventh aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line such that a ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 1.88, where ΔT is a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin.

[0024] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT at cruise to ΔT at takeoff is as follows: a) is between 0.48 and 1.71; b) is between 0.48 and 1.70 (where, for example, ΔT at take-off is ΔT at the take-off of the aircraft while it is on the ground); c) is between 0.48 and 1.55 (where, for example, ΔT at take-off is ΔT at the take-off of the aircraft while it is on the ground); d) is between 0.48 and 1.88 (where, for example, ΔT at takeoff is ΔT at the takeoff of the aircraft while it is in the air); or e) is between 0.48 and 1.71 (where ΔT at takeoff is ΔT at the takeoff of the aircraft while it is in the air).

[0025] According to an eighth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line such that a ratio of ΔT at cruise to ΔT at takeoff is between 0.48 and 1.88, where ΔT is a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin.

[0026] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of ΔT at cruise to ΔT at takeoff is as follows. a) is between 0.48 and 1.71; b) is between 0.48 and 1.70 (where, for example, ΔT at take-off is ΔT at the take-off of the aircraft while it is on the ground); c) is between 0.48 and 1.55 (where, for example, ΔT at take-off is ΔT at the take-off of the aircraft while it is on the ground); d) is between 0.48 and 1.88 (where, for example, ΔT at takeoff is ΔT at the takeoff of the aircraft while it is in the air); or e) is between 0.48 and 1.71 (where ΔT at takeoff is ΔT at the takeoff of the aircraft while it is in the air).

[0027] According to a ninth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line, the modulator valve being arranged to initiate the return of the fuel to the fuel tank when the fuel that has flowed through the heat exchanger has a temperature of at least 120°C.

[0028] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when the fuel flowing through the heat exchanger has a temperature of at least 140 °C.

[0029] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when the fuel flowing through the heat exchanger has a temperature between 120°C and 180°C, and preferably has a temperature between 140°C and 180°C.

[0030] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when: (i) an indication of an operating condition is provided; and ii) the fuel flowing through the heat exchanger has a temperature of at least 120 °C, preferably at least 140 °C.

[0031] The operating condition may be one or more of the following: a proportion of Sustainable Aviation Fuel (SAF) in the fuel, a thermal stability of the fuel, a degree of coking of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0032] The gas turbine engine may further include a sensor configured to detect one or more operating conditions.

[0033] According to a tenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line; wherein the method comprises initiating the return of fuel to the fuel tank using the modulator valve when the fuel flowing through the heat exchanger has a temperature of at least 120°C.

[0034] The method may include initiating the return of fuel to the fuel tank using the modulator valve when the fuel flowing through the heat exchanger has a temperature of at least 140°C.

[0035] The method may include initiating the return of fuel to the fuel tank using the modulator valve when the fuel flowing through the heat exchanger has a temperature between 120°C and 180°C, and preferably a temperature between 140°C and 180°C.

[0036] The method may include initiating the return of fuel to the fuel tank using the modulator valve when: (i) an indication of an operating condition is provided; and ii) the fuel flowing through the heat exchanger has a temperature of at least 120 °C, preferably at least 140 °C.

[0037] The operating condition may be one or more of the following: a proportion of Sustainable Aviation Fuel (SAF) in the fuel, a thermal stability of the fuel, a degree of coking of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0038] 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.

[0039] According to an eleventh aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve that modulates the fuel flow along the fuel return line so that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0040] The "mass of fuel returned to the fuel tank" and the "mass of fuel delivered to the combustion chamber" can be expressed as the fuel flow rate in terms of the mass of fuel flow per unit time (e.g., in kg / s). By "mass of fuel delivered to the combustion chamber," in the eleventh and twelfth aspects, we can mean the "mass of fuel burned," e.g., the amount of fuel burned by the combustion chamber in units of the fuel flow rate per unit time.

[0041] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0042] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0043] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0044] According to a twelfth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line; wherein the method comprises modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the fuel tank to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0045] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0046] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0047] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0048] According to a thirteenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line, the modulator valve being arranged to prevent the return of fuel having a temperature of 180 °C or higher to the fuel tank.

[0049] The modulator valve may be arranged to prevent the return of fuel to the fuel tank when the fuel temperature in the fuel tank reaches a predetermined upper threshold temperature.

[0050] 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.

[0051] The modulator valve may be arranged to modulate a fuel flow along the fuel return line based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a fuel temperature in the fuel tank, and an amount of fuel remaining in the tank.

[0052] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that an equilibrium temperature of the fuel in the fuel tank, after being mixed with fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature.

[0053] According to a fourteenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least a portion of the fuel that has flowed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the fuel flow along the fuel return line; wherein the method comprises preventing the return of fuel having a temperature of 180°C or higher to the fuel tank using the modulator valve.

[0054] The method may include preventing the return of fuel to the fuel tank using the modulator valve when the temperature of the fuel in the fuel tank reaches a predetermined upper threshold temperature.

[0055] 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.

[0056] The method may include modulating a fuel flow along the fuel return line using the modulator valve based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0057] The method may include modulating a fuel flow along the fuel return line using the modulator valve such that an equilibrium temperature of the fuel in the fuel tank after being mixed with fuel returned to the tank along the fuel return line does not exceed the predetermined upper threshold temperature.

[0058] According to a fifteenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger arranged to receive fuel and transfer heat from the oil to the fuel to raise the fuel temperature to at least 120°C upon entering the combustion chamber; a fuel recirculation line arranged to recirculate at least some fuel in 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 a modulator valve arranged to modulate the fuel flow along the fuel recirculation line.

[0059] The fuel recirculation line may be arranged 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.

[0060] The gas turbine engine may further include a fuel pump in the fuel flow path. The fuel recirculation line may be arranged to recirculate fuel from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

[0061] The fuel pump may be arranged in the fuel flow path downstream of the fuel-oil heat exchanger.

[0062] The fuel recirculation line may be arranged to supply fuel to one or more additional aircraft and / or engine mechanisms.

[0063] The one or more additional aircraft and / or engine mechanisms may be or include one or more of the following: engine nacelle anti-icing system, actuators, vent valves, thermal management modulation valves (for example, for the engine and / or a generator system and a turbine casing cooling system).

[0064] The fuel flow path may be a main 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 in the main fuel flow path.

[0065] The second point may be downstream of a low-pressure fuel pump that pumps fuel from the fuel tank to the gas turbine engine.

[0066] 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).

[0067] The engine may further include a secondary fuel-oil heat exchanger located in the main flowpath (e.g., in series with the main heat exchanger). The second point may be located downstream of the secondary heat exchanger.

[0068] Alternatively, the engine may further include a secondary fuel-oil heat exchanger having a flow line leading from the main flowpath to the secondary fuel-oil heat exchanger and a flow line rejoining the main flowpath from the secondary heat exchanger (e.g., running parallel to the main flowpath). 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 rejoining the main fuel flowpath from the secondary fuel-oil heat exchanger. The second point may be located upstream of the primary fuel-oil heat exchanger.

[0069] The fuel-oil heat exchanger may be arranged to transfer heat from the oil to the fuel to raise the fuel temperature to at least 140°C upon entering the combustion chamber (i.e., the heat exchanger may be arranged to transfer heat from the oil to the fuel to raise the fuel temperature to these temperatures before entering the combustion chamber).

[0070] The heat exchanger may be arranged to transfer heat from the oil to the fuel to increase the fuel temperature upon entering the combustion chamber to between 120 °C and 180 °C, preferably to between 140 °C and 180 °C.

[0071] The heat exchanger may be arranged to transfer heat from the oil to the fuel to increase the fuel temperature on entry into the combustion chamber 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, or to a fuel temperature in a range lying between any two of these values.

[0072] The modulator valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the recirculated fuel mass to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0073] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0074] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0075] The modulator valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0076] The gas turbine engine may further include at least one temperature sensor in the fuel flow path and / or in the fuel recirculation line.

[0077] The modulator valve may be arranged to modulate the fuel flow along the fuel recirculation line based at least in part on temperature data from at least one temperature sensor.

[0078] According to a sixteenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: a combustion chamber; a fuel-oil heat exchanger that receives fuel from a fuel tank and transfers heat from the oil to the fuel to raise the fuel temperature to at least 120°C upon entering the combustion chamber; a fuel recirculation line arranged to recirculate at least some fuel in 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 a modulator valve arranged to modulate the fuel flow along the fuel recirculation line. wherein the method comprises modulating the fuel flow along the fuel recirculation line using the modulator valve.

[0079] The method may include modulating the fuel flow along the fuel recirculation line using the modulator valve from a first point downstream of the fuel-oil heat exchanger to a second point upstream of the fuel-oil heat exchanger.

[0080] The gas turbine engine may further include a fuel pump in the fuel flow path. The method may include modulating the fuel flow along the fuel recirculation line using the modulator valve from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

[0081] The fuel pump may be arranged in the fuel flow path downstream of the fuel-oil heat exchanger.

[0082] The method may include supplying fuel to one or more additional aircraft and / or engine mechanisms via the fuel recirculation line.

[0083] The one or more additional aircraft and / or engine mechanisms may be or include one or more of the following: engine nacelle anti-icing system, actuators, vent valves, thermal management modulation valves (for example, for the engine and / or a generator system and a turbine casing cooling system).

[0084] The fuel flow path may be a main fuel 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 in the main fuel flow path.

[0085] The second point may be downstream of a low-pressure fuel pump that pumps fuel from the fuel tank to the gas turbine engine.

[0086] 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).

[0087] The engine may further include a secondary fuel-oil heat exchanger located in the main flowpath (e.g., in series with the main heat exchanger). The second point may be located downstream of the secondary heat exchanger.

[0088] Alternatively, the engine may further include a secondary fuel-oil heat exchanger having a flow line leading from the main flowpath to the secondary fuel-oil heat exchanger and a flow line rejoining the main flowpath from the secondary heat exchanger (e.g., running parallel to the main flowpath). 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 rejoining the main fuel flowpath from the secondary fuel-oil heat exchanger. The second point may be located upstream of the primary fuel-oil heat exchanger.

[0089] The method may include transferring heat from the second oil to the fuel (directly or indirectly) to assist in raising the fuel temperature to an average of at least 140°C upon entry into the combustion chamber under cruise conditions.

[0090] The method may include transferring heat from the oil to the fuel to increase the fuel temperature to 120°C to 180°C upon entering the combustion chamber.

[0091] The method may include transferring heat from the oil to the fuel to increase the fuel temperature to 140°C to 180°C upon entering the combustion chamber.

[0092] The heat exchanger may be arranged to transfer heat from the oil to the fuel to increase the fuel temperature on entry into the combustion chamber 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, or to a fuel temperature in a range lying between any two of these values.

[0093] The method may include modulating the fuel flow along the recirculation line using the modulator valve such that a ratio of the recirculated fuel mass to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0094] The method may include modulating the fuel flow along the recirculation line using the modulator valve such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0095] The method may include modulating the fuel flow along the recirculation line using the modulator valve such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0096] The method may include modulating the fuel flow along the recirculation line using the modulator valve such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0097] The gas turbine engine may further include at least one temperature sensor in the fuel flow path and / or in the fuel recirculation line.

[0098] The method may include modulating fuel flow along the fuel recirculation line using the modulator valve based at least in part on temperature data from the at least one temperature sensor.

[0099] 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 to the fuel tank using the modulation valve of any aspect defined above or elsewhere herein, as well as recirculated as defined in the fifteenth and sixteenth aspects. In the fifth, seventh, ninth, eleventh and thirteenth aspects:

[0100] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin, is less than 0.56 (e.g., under cruise conditions).

[0101] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is less than 0.53 (e.g., under cruise conditions).

[0102] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 0.56 (e.g., under cruise conditions).

[0103] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 0.53 (e.g., under cruise conditions).

[0104] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 1.00 (e.g., under cruise conditions).

[0105] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, (e.g., under cruise conditions) is as follows: a) is between 0.48 and 0.90; b) is between 0.48 and 0.82; c) is between 0.56 and 1.00; d) is between 0.56 and 0.82; or e) is between 0.56 and 0.75. In the sixth, eighth, tenth, twelfth and fourteenth aspects:

[0106] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of a temperature of the fuel in the fuel tank, in Kelvin, to a temperature of the fuel delivered to the combustion chamber, in Kelvin, is less than 0.56 (e.g., under cruise conditions).

[0107] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is less than 0.53 (e.g., under cruise conditions).

[0108] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 0.56 (e.g., under cruise conditions).

[0109] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 0.53 (e.g., under cruise conditions).

[0110] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, is between 0.48 and 1.00 (e.g., under cruise conditions).

[0111] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the temperature of the fuel in the fuel tank, in Kelvin, to the temperature of the fuel delivered to the combustion chamber, in Kelvin, (e.g., under cruise conditions) is as follows: a) is between 0.48 and 0.90; b) is between 0.48 and 0.82; c) is between 0.56 and 1.00; d) is between 0.56 and 0.82; or e) is between 0.56 and 0.75. In the first, third, fifth, seventh, eleventh and thirteenth aspects:

[0112] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when the fuel flowing through the heat exchanger has a temperature of at least 120°C and preferably at least 140°C.

[0113] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when the fuel flowing through the heat exchanger has a temperature between 120°C and 180°C, and preferably has a temperature between 140°C and 180°C.

[0114] The modulator valve may be arranged to initiate the return of fuel to the fuel tank when: (i) an indication of an operating condition is provided; and ii) the fuel flowing through the heat exchanger has a temperature of at least 120 °C, preferably at least 140 °C.

[0115] The operating condition may be one or more of the following: a proportion of Sustainable Aviation Fuel (SAF) in the fuel, a thermal stability of the fuel, a degree of coking of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0116] The gas turbine engine may further include a sensor configured to detect one or more operating conditions. In the second, fourth, sixth, eighth, twelfth and fourteenth aspects:

[0117] The method may include initiating the return of fuel to the fuel tank using the modulator valve when the fuel flowing through the heat exchanger has a temperature of at least 120°C, and preferably at least 140°C.

[0118] The method may include initiating the return of fuel to the fuel tank using the modulator valve when the fuel flowing through the heat exchanger has a temperature between 120°C and 180°C, preferably a temperature between 140°C and 180°C.

[0119] The method may include initiating the return of fuel to the fuel tank using the modulator valve when: (i) an indication of an operating condition is provided; and ii) the fuel flowing through the heat exchanger has a temperature of at least 120 °C, preferably at least 140 °C.

[0120] The operating condition may be one or more of the following: a proportion of Sustainable Aviation Fuel (SAF) in the fuel, a thermal stability of the fuel, a degree of coking of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0121] 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. In the first, third, fifth, seventh, ninth and thirteenth aspects

[0122] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of fuel returned to the tank to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0123] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0124] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0125] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions. In the second, fourth, sixth, eighth, tenth and fourteenth aspects:

[0126] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that the ratio of the fuel mass returned to the fuel tank to the fuel mass delivered to the combustion chamber is between 0 and 9 under cruise conditions.

[0127] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is between 2.3 and 9 under cruise conditions.

[0128] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 9 under cruise conditions.

[0129] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that a ratio of the fuel mass returned to the tank to the fuel mass delivered to the combustion chamber is greater than 4 and less than or equal to 4.9 under cruise conditions. In the first, third, fifth, seventh, ninth and eleventh aspects:

[0130] The modulator valve may be arranged to prevent the return of fuel having a temperature of 180 °C or higher to the fuel tank.

[0131] The modulator valve may be arranged to prevent the return of fuel to the fuel tank when the fuel temperature in the fuel tank reaches a predetermined upper threshold temperature.

[0132] 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.

[0133] The modulator valve may be arranged to modulate a fuel flow along the fuel return line based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a fuel temperature in the fuel tank, and an amount of fuel remaining in the tank.

[0134] The modulator valve may be arranged to modulate the fuel flow along the fuel return line such that an equilibrium temperature of the fuel in the fuel tank, after being mixed with fuel returned to the tank along the fuel return line, does not exceed the predetermined upper threshold temperature. In the second, fourth, sixth, eighth, tenth and twelfth aspects:

[0135] The method may include preventing the return of fuel having a temperature of 180°C or higher to the fuel tank by using the modulator valve.

[0136] The method may include preventing the return of fuel to the fuel tank using the modulator valve when the temperature of the fuel in the fuel tank reaches a predetermined upper threshold temperature.

[0137] 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.

[0138] The method may include modulating a fuel flow along the fuel return line using the modulator valve based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0139] The method may include modulating a fuel flow along the fuel return line using the modulator valve such that an equilibrium temperature of the fuel in the fuel tank after being mixed with fuel returned to the tank along the fuel return line does not exceed the predetermined upper threshold temperature. In the first, second, third and fourth aspects:

[0140] A temperature of the fuel fed to the combustion chamber (i.e., upon entry into the combustion chamber) may be a temperature of the fuel fed to the combustion chamber under cruise flight conditions. Fuel temperatures upon entry into the combustion chamber under cruise flight conditions may be defined as an average over at least 5 minutes, 10 minutes, or 30 minutes under steady cruise flight conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in the temperature of the fuel during operation, often increasing the temperature. Each fluctuation must not last more than 5 minutes.

[0141] The previous paragraph above may also be relevant for other references to the temperature of the fuel entering the combustion chamber. In the first, third, fifth, seventh, ninth, eleventh, thirteenth and fifteenth aspects:

[0142] The heat exchanger may be arranged to transfer heat from the oil to the fuel, thereby increasing the fuel temperature upon entry into the combustion chamber to at least 120°C, and preferably to 140°C (i.e., the heat exchanger may be arranged to transfer heat from the oil to the fuel, thereby increasing the temperature of the fuel so that the fuel is at that temperature upon entry into the combustion chamber).

[0143] The heat exchanger may be arranged to transfer heat from the oil to the fuel to increase the fuel temperature upon entering the combustion chamber to between 120 °C and 180 °C, and preferably to between 140 °C and 180 °C.

[0144] The heat exchanger may be arranged to transfer heat from the oil to the fuel to increase the fuel temperature on entry into the combustion chamber 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, or to a fuel temperature in a range lying between any two of these values. In the second, fourth, sixth, eighth, tenth and twelfth, fourteenth and sixteenth aspects:

[0145] The method may include transferring heat from the oil to the fuel using the fuel-oil heat exchanger to increase the fuel temperature upon entering the combustion chamber to at least 120°C, and preferably to at least 140°C.

[0146] The method may include transferring heat from the oil to the fuel using the fuel-oil heat exchanger to increase the fuel temperature upon entering the combustion chamber to between 120°C and 180°C, and preferably to between 140°C and 180°C.

[0147] The method may include transferring heat from the oil to the fuel to increase the fuel temperature upon entry into the combustion chamber 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, or to a fuel temperature in a range between any two of these values. In the first, third, fifth and seventh aspects:

[0148] 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;

[0149] The first temperature sensor may be located in the fuel return line.

[0150] The modulator valve may be arranged to modulate the fuel flow along the fuel return line based on temperature data from the first temperature sensor and the second temperature sensor. In the ninth aspect:

[0151] The gas turbine engine may further include at least one temperature sensor located downstream of the heat exchanger.

[0152] The modulator valve may be arranged to initiate the return of fuel to the fuel tank based at least in part on temperature data from at least one temperature sensor located downstream of the heat exchanger. In the eleventh and thirteenth aspects:

[0153] The gas turbine engine may further include 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.

[0154] The at least one temperature sensor located downstream of the heat exchanger may be located in the fuel return line.

[0155] The modulator valve may be arranged to modulate the fuel flow 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. In the second, fourth, sixth and eighth aspects:

[0156] 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;

[0157] The first temperature sensor may be located in the fuel return line.

[0158] The method may include modulating the fuel flow along the fuel return line based on temperature data from the first temperature sensor and the second temperature sensor. In the tenth aspect:

[0159] The gas turbine engine may further include at least one temperature sensor located downstream of the heat exchanger.

[0160] The method may include initiating the return of fuel to the fuel tank using the modulator valve based at least in part on temperature data from at least one temperature sensor located downstream of the heat exchanger. In the twelfth and fourteenth aspects:

[0161] The gas turbine engine may further include 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.

[0162] The at least one temperature sensor located downstream of the heat exchanger may be located in the fuel return line.

[0163] The method may include modulating the fuel flow along the fuel return line using the modulator 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. In the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteenth and fourteenth aspects:

[0164] The temperature of the fuel in the fuel tank can range from -54 °C to 100 °C.

[0165] The temperature of the fuel in the fuel tank can range from -54 °C to 65 °C.

[0166] The temperature of the fuel in the fuel tank can be between -54 °C and 55 °C.

[0167] The temperature of the fuel in the fuel tank can range from -54 °C to 5 °C.

[0168] The temperature of the fuel in the fuel tank can 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 range between two of these values. In the first, third, fifth, seventh, ninth, eleventh and thirteenth aspects:

[0169] The modulator valve may be arranged to modulate the fuel flow along the fuel return line based at least in part on the amount of fuel remaining in the fuel tank. The amount of fuel remaining in the fuel tank may be a fuel mass remaining in the fuel tank.

[0170] The modulator valve may be arranged to modulate the fuel flow along the fuel return line based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a fuel temperature in the fuel tank, and an amount of fuel remaining in the fuel tank.

[0171] The modulator valve may be arranged to modulate the fuel flow 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. In the second, fourth, sixth, eighth, tenth, twelfth and fourteenth aspects:

[0172] The method may include modulating fuel flow along the fuel return line using the modulator valve 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.

[0173] The method may include modulating fuel flow along the fuel return line using the modulator valve based at least in part on a temperature of the fuel that has flowed through the heat exchanger, a temperature of the fuel in the fuel tank, and an amount of fuel remaining in the fuel tank.

[0174] The method may include modulating the fuel flow along the fuel return line using the modulator valve such that an equilibrium temperature of the fuel in the fuel tank after returning the fuel to the fuel tank along the fuel return line does not exceed a predetermined upper threshold temperature for the fuel in the fuel tank.

[0175] As indicated elsewhere herein, the present disclosure may be applicable to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open-rotor gas turbine engine (in which the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Each such engine may or may not be provided with an afterburner. Such a gas turbine engine may be configured, for example, for land-based or marine power generation applications.

[0176] A gas turbine engine according to the present disclosure may include an engine core including 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 (with fan blades). Such a fan may be disposed upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan disposed downstream of the engine core, for example, if the gas turbine engine has an open rotor or is a turboprop engine (in which case the fan may be referred to as a propeller).

[0177] If the gas turbine engine has an open rotor or is a turboprop engine, the gas turbine engine may comprise two counter-rotating propeller stages connected by a shaft to and driven by a free power turbine. The propellers may rotate in opposite directions, so that one rotates clockwise and the other counterclockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may comprise a propeller stage and a guide vane stage configured downstream of the propeller stage. The guide vane stage may have a variable pitch. Accordingly, high-pressure, intermediate-pressure, and free power turbines may each drive high- and intermediate-pressure compressors and propellers through suitable connecting shafts. The propellers may thus provide the majority of the propulsive thrust.

[0178] If the gas turbine engine has an open rotor or is a turboprop engine, one or more of the propeller stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0179] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct-drive turbofan engine in which the fan is connected directly to the fan drive turbine via a core shaft, for example, without a gearbox. In such a direct-drive turbofan engine, the fan may be arranged to rotate at the same speed as the fan drive turbine. For example only, 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 arranged to rotate at a higher speed than the first core shaft. In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0180] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine includes a fan driven by a gearbox. Accordingly, such a gas turbine engine may include a gearbox that receives input from the core shaft and delivers output to the fan such that the fan is driven at a lower speed than the core shaft. The drive to the gearbox may be directly from the core shaft or indirectly from the core shaft, for example, via a spur gear shaft and / or a gear. The core shaft may rigidly connect the turbine and compressor such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0181] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any number of shafts connecting turbines and compressors, for example, one, two, or three shafts. For 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 arranged to rotate at a higher speed than the first core shaft.

[0182] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., directly, e.g., via a generally annular channel).

[0183] The gearbox may be arranged to be driven by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., only by the first core shaft and not by the second core shaft in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example, the first and / or second shafts in the example above.

[0184] The gearbox may be a reduction gearbox (in that the output to the fan has a lower speed than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary gearbox" or a "star gearbox," as described in more detail elsewhere herein. Such a gearbox may be single-stage. Alternatively, such a gearbox may be a link gearbox, for example, a link planetary gearbox (which may have the input at the sun gear and the output at the ring gear and may thus be referred to as a "link star" gearbox), which has, for example, two reduction stages.

[0185] The transmission may have any desired reduction ratio (defined as the speed of the input shaft divided by the speed of the output shaft), for example greater than 2.5, for example in the range 3 to 4.2 or 3.2 to 3.8, for example in the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may, for example, be between any two of the values ​​mentioned in the preceding sentence. By way of example only, the transmission may be a "star gear" with a gear ratio in the range 3.1 or 3.2 to 3.8. Further by way of example only, the transmission may be a "star gear" with a gear ratio in the range 3.0 to 3.1. Further, purely by way of example, the transmission may be a "planetary gear" with a gear ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside these ranges.

[0186] In any gas turbine engine described and / or claimed herein, fuel of a given composition or mixture is provided to a combustor, which may be provided downstream of the fan and compressor(s) with respect to the flowpath (e.g., axially downstream). For example, the combustor may be located directly downstream of the second compressor (e.g., at its exit), where a second compressor is provided. As another example, flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0187] The or each compressor (for example, the first compressor and the second compressor, as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (since their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct-drive turbofan 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 another example, the gas turbine engine may be a "geared" gas turbine engine (where the fan is driven by a first core shaft via a reduction gearbox) that includes 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may include 3 or 4 stages in the first (or "low-pressure") compressor and 8 or 9 stages in the second (or "high-pressure") compressor. As another example, the gas turbine engine may be a "geared" gas turbine engine with 4 stages in the first (or "low-pressure") compressor and 10 stages in the second (or "high-pressure") compressor.

[0188] The or each turbine (for example, the first turbine and the second turbine as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise one row of rotor blades and one row of stator blades, or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from one another. The second (or "high pressure") turbine may comprise two stages in any arrangement (for example, regardless of whether it is a geared or direct drive engine). The gas turbine engine may be a direct drive turbofan engine comprising a first (or "low pressure") turbine with five, six, or seven stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine having a first (or "low pressure") turbine with three or four stages.

[0189] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-swept or 0% span location to a tip at a 100% span location. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or on the order 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 radius of the fan blade at the hub to the radius of the fan blade 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, in the range 0.28 to 0.32 or 0.29 to 0.30. These ratios may be generally referred to as the hub-to-tip ratio.The hub radius and the tip radius can both be measured at the leading edge (or axially forwardmost part) of the blade. The hub-to-tip ratio, of course, refers to the gas-swept section of the fan blade, i.e., the section radially outward from any platform.

[0190] The radius of the fan can be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The fan diameter (which can be simply twice the fan radius) can be larger than (or on the order 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 (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches) or 420 cm (about 165 inches). The fan diameter can be in an inclusive range limited by any two of the values ​​in the previous sentence (i.e., the values ​​can be upper or lower limits), for example, in the range from 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. Purely as a non-limiting example, the fan diameter can be in the range from 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.

[0191] The speed of the fan may vary during use. Generally, the speed is lower for fans with a larger diameter. Purely as a non-limiting example, the fan speed under cruise conditions may be less than 3500 rpm, for example, less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Purely as a further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm. Purely as a further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 2500 to 2800 rpm.By way of further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. By way of further non-limiting example, the fan speed under cruise conditions for a direct drive engine with a fan diameter in the range of 190 cm to 200 cm may be in the range of 3600 to 3900 rpm. By way of further non-limiting example, the fan speed under cruise conditions for a direct drive engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

[0192] During operation of the gas turbine engine, the fan (with associated fan blades) rotates around a rotational axis. This rotation causes the tip of the fan blade to rotate at a speed U SpitzeThe work performed by the fan blades in the flow leads to an enthalpy increase dH of the flow. A fan peak load can be expressed as dH / U Spitze 2 where dH is the enthalpy rise (for example, the mean 1-D enthalpy rise) across the fan and U Spitzeis the (translational) fan tip velocity, for example, at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip loading under cruise conditions can be greater than (or on the order 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 fan tip loading may be in an inclusive range bounded by any two of the values ​​in the previous sentence (i.e., the values ​​may form upper or lower limits), for example in the range 0.28 to 0.31 or 0.29 to 0.3 (for example, for a geared gas turbine engine).

[0193] Gas turbine engines according to the present disclosure may have a desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass channel to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio under cruise conditions may be greater than (or of the order of) one 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 bypass ratio under cruise conditions 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 12 to 16, 13 to 15, or 13 to 14.By way of non-limiting example only, the bypass ratio under cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. By way of further non-limiting example only, the bypass ratio of a geared gas turbine engine according to the present disclosure under cruise conditions may be in the range of 12:1 to 15:1. The bypass duct may be substantially annular. The bypass duct may be located radially outward from the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0194] The total pressure ratio (OPR) of a gas turbine engine, as described and / or claimed herein, may be defined as the ratio of the ram pressure at the exit of the maximum pressure compressor (before entering the combustor) to the ram pressure upstream of the fan. As a non-limiting example, the total pressure ratio of a gas turbine engine, as described and / or claimed herein, under cruise conditions may be greater than (or on the order of) one of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 50 to 70.By way of non-limiting example only, the cruise overall pressure ratio 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. By way of non-limiting example only, the cruise overall pressure ratio 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. By way of non-limiting example only, the cruise overall pressure ratio 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. By way of non-limiting example only, the cruise overall pressure ratio 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.

[0195] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, the specific thrust for a given thrust condition may depend on the specific composition of fuel provided to the combustion chamber. Under cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) one of the following: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s or 80 Nkg -1 s. The specific thrust may lie within an inclusive range limited by any two of the values ​​in the previous sentence (ie the values ​​may form upper or lower limits), for example in the range of 80 Nkg -1 s up to 100 Nkg -1 s or 85 Nkg -1 s up to 95 Nkg-1 s. Such engines can be particularly efficient compared to conventional gas turbine engines. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 90 Nkg -1 s up to 95 Nkg -1 s. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 80 Nkg -1 s up to 90 Nkg -1 s. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 70 Nkg -1 s up to 90 Nkg -1 s. Purely as a non-limiting example, the specific thrust of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 90 Nkg-1 s up to 120 Nkg -1 s be.

[0196] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Purely as a non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least one of (or on the order 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 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). Purely as a non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN.By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. By way of non-limiting example only, the maximum thrust of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN.The above thrust may be the maximum net thrust at normal atmospheric conditions at sea level at +15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C) with static engine.

[0197] In use, the temperature of the flow at the inlet to the high-pressure turbine may be particularly high. This temperature, which may be referred to as the TET, may be measured at the exit to the combustor, for example, immediately upstream of the first turbine guide vane, which may itself be referred to as the nozzle guide vane. In some examples, the TET for a given thrust condition may depend on the specific composition of fuel provided to the combustor. Under cruise conditions, the TET may be at least one of (or on the order of) the following: 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 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, 1530 K to 1600 K.Thus, purely as a non-limiting example, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 1540 K to 1600 K. Purely as a non-limiting example, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 1590 K to 1650 K. Purely as a non-limiting example, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1600 K to 1660 K. Purely as a non-limiting example, the TET of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm under cruise conditions may be in the range of 1590 K to 1650 K.Purely as a non-limiting example, the TET under cruise conditions of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1570 K to 1630 K.

[0198] The maximum TET in use of the engine may, for example, be at least one of (or on the order of) the following: 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 1800 K to 1950 K or 1900 K to 2000 K. Thus, purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range 200 cm to 210 cm may be in the range 1890 K to 1960 K. Purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K.Purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Purely as a non-limiting example, the maximum TET of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Purely as a non-limiting example, the maximum TET of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may, for example, occur at a high thrust condition, for example at a maximum lift-off (MTO) condition.

[0199] A fan blade and / or a vane portion of a fan blade as described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or vane may be made at least partially from a composite material, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As another example, at least a portion of the fan blade and / or vane may be made at least partially from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy), or a steel-based material. The fan blade may include at least two regions made using different materials.For example, the fan blade may include a protective leading edge, which may be fabricated using a material better able to withstand impact (e.g., from birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be fabricated using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may include a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0200] A fan as described and / or claimed herein may comprise a central portion from which the fan blades may extend, for example, in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fastener engageable with a corresponding slot in the hub (or disc). Such a fastener may, by way of example only, be in the form of a dovetail that is insertable and / or snap-fittable into a corresponding slot 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 blade disc or blade ring. Any suitable method may be used to manufacture such a blade disc or blade ring.For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be attached to the hub / disk by welding, such as linear friction welding.

[0201] The gas turbine engines described and / or claimed herein may or may not be equipped with a variable area nozzle (VAN). Such a variable area nozzle may enable 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.

[0202] The fan of a gas turbine as described and / or claimed herein may have any number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. If the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. If the fan blades have a metallic body (for example, an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0203] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more parts thereof) have their conventional meaning and are readily understood by those skilled in the art. Thus, those skilled in the art would readily recognize that, for a given gas turbine engine for an aircraft, each term refers to all or one or more parts of an operating phase of the engine within a given application of an aircraft to which the gas turbine engine is to be attached.

[0204] In this regard, ground idle may refer to a phase of engine operation in which the aircraft is stationary and in contact with the ground, but with a requirement that the engine be running. During idle, the engine may produce between 3% and 9% of the engine's available thrust. In further non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In further non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxi may refer to a phase of engine operation in which the aircraft is propelled along the ground by the thrust generated by the engine. During taxi, the engine may produce between 5% and 15% of the available thrust. In further non-limiting examples, the engine may produce between 6% and 12% of the available thrust.In further non-limiting examples, the engine may produce between 7% and 10% of the available thrust. Takeoff may refer to a phase of engine operation in which the aircraft is propelled by the thrust generated by the engine. In an initial stage within the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. In a later stage within 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 further non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In further non-limiting examples, the engine may produce 100% of the available thrust.

[0205] Climb may refer to a phase of engine operation in which the aircraft is propelled by the thrust generated by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In further non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In further non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and the achievement of cruise conditions, with the achievement of cruise conditions thus defining the beginning of the cruise phase or portion thereof of the aircraft flight.Additionally or alternatively, climb may refer to a nominal point or to one or more nominal periods during an aircraft flight cycle between takeoff and landing when a relative increase in altitude is required, which may require additional engine thrust.

[0206] As used herein, cruise conditions, which define the cruise phase (or portion thereof) of flight, have the conventional meaning and are 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 engine operating point at mid-cruise of a given mission (what may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is to be mounted. In this regard, mid-cruise may be the point in an aircraft flight cycle at which 50% of the total fuel burned between the peak of climb and the beginning of descent has been burned (which may be approximated by the midpoint—in time and / or distance—between the peak of climb and the beginning of descent).The cruise conditions may thus define an operating point, phase or portion thereof of the flight providing thrust that would 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 an at least substantially constant Mach number) of an aircraft to which it is to be mounted, taking into account the number of engines provided for that aircraft.For example, if an engine is designed to be fitted to an aircraft having two engines of the same type, the engine will provide, under cruise conditions, half the total thrust that would be required for steady-state, or at least substantially steady-state, operation of that aircraft at mid-cruise.

[0207] In other words, the cruise conditions for a given aircraft gas turbine engine can be defined as the engine operating point that provides a specified thrust (required to provide, in combination with any other engines on the aircraft, steady-state, or at least substantially steady-state, operation of the aircraft to which it is to be fitted, at a given mid-cruise Mach number) under mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at cruise altitude). For any given aircraft gas turbine engine, the mid-cruise thrust, atmospheric conditions, and Mach number can be known, and thus the engine operating point under cruise conditions is clearly defined.

[0208] By way of example only, forward speed under cruise conditions may be any point in the range of Mach 0.7 to 0.9, for example, 0.75 to 0.85, for example, 0.76 to 0.84, for example, 0.77 to 0.83, for example, 0.78 to 0.82, for example, 0.79 to 0.81, for example, on the order of Mach 0.8, on the order of Mach 0.85, or in the range of 0.8 to 0.85. Any individual speed within these ranges may be part of the cruise conditions. For some aircraft, cruise conditions may be outside these ranges, for example, below Mach 0.7 or above Mach 0.9.

[0209] By way of example only, the cruise flight conditions may correspond to standard atmosphere 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 (about 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (about 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range of 10,900 m to 11,100 m, for example in the order of 11,000 m. Cruise conditions can be equivalent to standard atmospheric conditions at any given altitude in these areas.

[0210] For 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 can provide a known required net thrust level. The known required net thrust level naturally depends on the engine and its intended use and may, for example, be a value in the range of 20 kN to 40 kN.

[0211] Further, purely by way of example, 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 naturally depends on the engine and its intended use and may, for example, be a value in the range of 35 kN to 65 kN.

[0212] In use, a gas turbine engine as described and / or claimed herein may operate under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., the mid-cruise conditions) of an aircraft on which at least one (e.g., two or four) gas turbine engines may be mounted to provide propulsive thrust.

[0213] Furthermore, one of ordinary skill in the art would readily recognize that either or both of descent and approach refer to an operational phase within an aircraft flight cycle between cruise and landing of the aircraft, with approach forming part of the landing and takeoff (LTO) phase. During either or both of descent and approach, the engine may produce between 0% and 50% of available thrust. In further non-limiting examples, the engine may produce between 25% and 40% of available thrust. In further non-limiting examples, the engine may produce between 30% and 35% of available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing at which a relative decrease in altitude is required, which may require a reduced thrust requirement from the engine.

[0214] According to one aspect, an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. The aircraft according to this aspect is the aircraft for which the gas turbine engine is designed to be mounted. Accordingly, the cruise flight conditions according to this aspect may correspond to an operating point, phase, or portion thereof of the aircraft's flight, as defined elsewhere herein.

[0215] According to one aspect, a method of operating a gas turbine engine as described and / or claimed herein is provided. Operation may be under suitable conditions as may be defined elsewhere herein (e.g., with respect to thrust, atmospheric conditions, and Mach number).

[0216] According to one aspect, a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. Operation according to this aspect may include (or be) operation under suitable conditions, for example, at mid-cruise of the aircraft, as defined elsewhere herein.

[0217] Those skilled in the art will understand that, except in cases of mutual exclusion, a feature or parameter described with respect to one of the above aspects may be applied to any other aspect. Furthermore, except in cases of mutual exclusion, any feature or parameter included or described herein may be applied to any aspect and / or combined with any other feature or parameter described herein.

[0218] Except in cases of mutual exclusion, any parameter or value included or described herein may be applied to and / or combined with one or more other parameters and / or values ​​included or described herein. For example, a first parameter or value included or described herein (e.g., Parameter A) may be applied to and / or combined with one or more other parameters and / or values ​​included or described herein (e.g., one or more of Parameter B; Parameter C; and Parameter D, etc.) to express a product of their relationship. For example, one of skill in the art would understand that if Parameter A is disclosed separately from Parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or any other such application, combination, or function of Parameter A relative to Parameter B, as appropriate.

[0219] Embodiments will now be described by way of example only with reference to the figures in which: Fig. Figure 1 is a cross-sectional side view of a gas turbine engine; Fig. Figure 2 is a close-up cross-sectional side view of an upstream section of a geared gas turbine engine; Fig. Figure 3 is a partially cutaway view of a gearbox for a gas turbine engine; Fig. Figure 4 is a close-up cross-sectional side view of an upstream section of a direct drive gas turbine engine; Fig. 5 is a schematic diagram of an exemplary fuel system including a fuel return line; Fig. 6 is a schematic diagram of an exemplary fuel system including a fuel recirculation line; Fig. 7 is a schematic illustration of another exemplary fuel system including a fuel return line; Fig. 8 to 13 illustrate an exemplary method for operating a gas turbine engine; Fig. 14 is a schematic diagram of an exemplary fuel system including a fuel recirculation line; Fig. 15 is a schematic diagram of an exemplary fuel system including a fuel recirculation line and a fuel return line; Fig. 16 is a schematic representation of an exemplary fuel system including a fuel recirculation line; and Fig. Figure 17 is an illustration of an aircraft having a propulsion system including two gas turbine engines.

[0220] Fig. 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air inlet 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow sequence, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to the low-pressure turbine 19 via a shaft 26 and an epicyclic gear train 30 and is driven thereby.

[0221] In use, the core air stream A is accelerated and compressed by the low-pressure compressor 14 and passed into the high-pressure compressor 15 where it is further compressed. The compressed air discharged from the high-pressure compressor 15 is passed 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 a combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through, 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 via a suitable connecting shaft 27.Fan 23 generally acts to apply increased pressure to the bypass airflow B flowing through bypass duct 22, so that bypass airflow B is expelled through bypass exhaust nozzle 18 to generally provide the majority of the propulsive thrust. Epicyclic gearing 30 is a reduction gear.

[0222] An exemplary arrangement for a geared turbofan 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 central pinion or sun gear 28 of the epicyclic gear assembly 30. Radially outward from and meshing with the sun gear 28 are a plurality of planet gears 32 coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to orbit synchronously about the sun gear 28 while enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via linkages 36 to drive its rotation about the engine axis 9. Radially outward from and meshing with the planet gears 32 is a ring gear or ring gear 38 coupled to a stationary support structure 24 via linkages 40.

[0223] It should be noted that the terms "low pressure turbine" and "low pressure compressor," as used herein, mean the lowest pressure turbine stages and the lowest pressure compressor stages, respectively (i.e., not including the fan 23) and / or the turbine and compressor stages connected by the lowest speed connecting shaft 26 in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some references, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor." Where such alternative nomenclature is used, the fan 23 may be referred to as a first or lowest pressure compression stage.

[0224] The epicyclic gear 30 is exemplary in Fig. 3 in more detail. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth on its circumference to mesh with the other gears. For clarity, Fig. 3, however, illustrates only exemplary portions of the teeth. Four planetary gears 32 are illustrated, although it will be apparent to one skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary gear train 30 generally include at least three planetary gears 32.

[0225] This is exemplified in the Fig. 2 and Fig. The epicyclic gear train 30 illustrated in Figure 3 is of the planetary type, as the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 being fixed. However, any other suitable type of epicyclic gear train 30 may be used. As another example, the epicyclic gear train 30 may be a star arrangement, in which the planet carrier 34 is held fixed and the ring gear (or ring gear) 38 is allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the train 30 may be a differential gear train, in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0226] It is understood that the Fig. 2 and Fig. 3 is merely exemplary and that various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement for housing the transmission 30 in the engine 10 and / or for connecting the transmission 30 to the engine 10 may be used. As another example, the connections (such as the linkages 36, 40 in the example of Fig. 2) between the transmission 30 and other parts of the engine 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 engine (for example, between the input and output shafts of the transmission and the fixed structures, such as the transmission housing) may be used, and the disclosure is not limited to the exemplary arrangement of Fig. 2. For example, if the gearbox 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of output and support rods and bearing positions would typically be different from that shown in Fig. 2 would deviate from the arrangement shown as an example.

[0227] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (for example, star or planetary gears), support structures, input and output shaft arrangement, and bearing positions.

[0228] Optionally, the gearbox can drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster compressor).

[0229] 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 compressors and / or turbines and / or an alternative number of connecting shafts. As another example, the Fig. 1 has a split-flow nozzle 18, 20, meaning that the flow through the bypass duct 22 has its own nozzle 18 separate from and located radially outward from 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. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

[0230] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a gearbox for the main shaft(s) and may instead be direct drive engines. A cross-sectional view of such an engine is shown in Fig. 4 shown.

[0231] With reference to Fig. 4, a gas turbine engine is indicated generally at 10 with a main axis of rotation 9. The engine 10 includes, in axial flow order, an air inlet 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine core 10 and defines both the inlet 12 and the exhaust nozzle 20.

[0232] In use, air entering inlet 12 is accelerated by fan 23 to create two airflows: a core airflow A and a bypass airflow B. Core airflow A flows into intermediate pressure compressor 14, and bypass airflow B passes through a bypass duct 22 to provide motive thrust. Intermediate pressure compressor 14 compresses airflow A before discharging the air to high pressure compressor 15, where further compression occurs.

[0233] 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 a combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through, 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 turbine 17, the intermediate-pressure turbine 19a, and the low-pressure turbine 19 each drive the high-pressure compressor 15, the intermediate-pressure compressor 14, and the fan 23, respectively, through a suitable connecting shaft.

[0234] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of connecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a gearbox provided in the driveline from a turbine to a compressor and / or a fan.

[0235] While the described example refers to a turbofan engine, the disclosure may be applied to any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0236] The geometry of the gas turbine engine 10 and components thereof is defined by a conventional axis system comprising an axial direction (aligned with the rotational axis 9), a radial direction (in the direction from bottom to top in Fig. 1) and a circumferential direction (perpendicular to the side in the view of Fig. 1). The axial, radial, and circumferential directions are perpendicular to each other.

[0237] The fuel F provided 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, the fuel F, when blended, mixed, or replaced with an alternative fuel, may comprise renewable hydrocarbons produced from biological or non-biological resources, also known as Sustainable Aviation Fuel (SAF). In each of the provided examples, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0238] For example, SAF is understood by those skilled in the art to mean a biofuel, a renewable aviation fuel, a renewable aviation fuel, an alternative fuel, or a biojet fuel produced from biological or non-biological resources. SAF is typically synthesized from carbon-containing gases extracted from the atmosphere and / or from industrial processes; or from a wide range of sustainable feedstocks such as waste oils and fats; municipal solid waste; cellulosic waste (such as corn stalks); cover crops such as camelina, carinata, and rockcress; non-biogenic alternative fuels; jatropha; halophytes, and algae, rather than from fossil hydrocarbons. SAF is understood to exclude fossil fuels.

[0239] The functional performance of a given fuel composition or blend of fuel F for use in a given application may be defined, at least in part, by the ability of the fuel to serve the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and emissions including gaseous and / or particulate matter. In this regard, particulate matter emissions may include soot particles produced during the combustion of the fuel F, also referred to as non-volatile particulate matter (nvPM). Any references herein to soot or smoke may equally apply to other types of particulate matter emissions known in the art.Gaseous emissions may include one or more of the following nitrogen oxides (NOx): carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO), including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOCs) produced during the combustion of the fuel F. All references herein to gaseous emissions may equally apply to other types of gaseous emissions known in the art.

[0240] A relatively higher specific energy (i.e., energy per unit mass), expressed as MJ / kg, can 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 as MJ / l, can at least partially reduce takeoff fuel volume, which may be particularly important for volume-limited or single-refueling military operations. A relatively higher thermal stability (i.e., inhibition of fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can enable reduced contrail formation while reducing the environmental impact of a given mission. Other fuel properties can also be central to functional performance. For example, a relatively lower freezing point (°C) can enable long-range missions to optimize flight profiles; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to fuels with high aromatic content; and maximum surface tension (mN / m) can ensure sufficient spray dispersion and atomization of the fuel.

[0241] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can affect the specific energy of a given fuel composition or blend. Fuels with higher ratios of hydrogen atoms to carbon atoms can have higher specific energies in the absence of bond strain. For example, fossil-based hydrocarbon fuels may comprise molecules of approximately 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.

[0242] 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 between 10% and 50% sustainable aviation fuel (with the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with further compositions pending approval. However, it is anticipated within 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 consisting of 100% SAF or to a fuel containing SAF, e.g., a SAF blend.

[0243] Sustainable aviation fuels may comprise one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and may be made, 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 comprise one or both of a lower aromatic and sulfur content relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise one or both of a higher isoalkane and cycloalkane content relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may comprise one or both of a density of between 90% and 98% of that of kerosene and a calorific value of between 101% and 105% of that of kerosene.

[0244] In some examples, the sustainable aviation fuel(s), or blend(s), provided to combustion equipment 16 may have a relatively lower content of aromatics and / or other non-paraffinic ingredients than kerosene. The sustainable aviation fuel may include an aromatics content of, for example, 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; e.g., 4%, 3%, 2%, 1%, or less than 1%; e.g., 0.75%, 0.5%, 0.25%, or less than 0.25%; e.g., 0.2%, 0.1%, or less than 0.1%; e.g., 0.01%, 0.001%, or 0%. The aroma content of the sustainable aviation fuel may be within an inclusive value or range limited by or within any two values ​​from the previous sentence (i.e. the values ​​may form upper or lower limits), e.g.13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005%, or 0%; or 0% to 0.75%, 0% to 0.5%, or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%; depending on one or more of preference, fuel stock or supplier, and variations in composition.

[0245] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide benefits 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). Accordingly, sustainable aviation fuels may result in one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs compared to fossil-based hydrocarbon fuels such as kerosene.

[0246] 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 schematically shown in Fig. 5. In the example shown, the engine 10 comprises a gearbox 30, as shown in the Fig. 1 to 3 (however, it may be used in conjunction with other gas turbine engines, including those with a direct drive architecture). The fuel system 1000 includes both the fuel delivery system (comprising the fuel tank 50 and pump 1002) that supplies the engine and the fuel management system 1500 (operating within the engine) of the aircraft. The fuel management system 1500 manages the fuel temperature and fuel flow, directing the fuel through one or more heat exchangers 1004, 1006 of the engine's heat exchange system.

[0247] In the described implementation, each engine 10 has 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, be a duplicate of several of the Fig. 5 for the other engine(s).

[0248] A low-pressure fuel supply pump 1002 pumps fuel from the fuel tank 50 to the gas turbine engine 10. The fuel then flows through a secondary fuel-oil heat exchanger 1004 and a primary fuel-oil heat exchanger 1006 before flowing 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 described as a primary fuel-oil heat exchanger because the oil flowing through it can be used to cool and lubricate the main gearbox 30 of the engine 10.The secondary fuel-oil heat exchanger 1004 may be described as a generator fuel-oil heat exchanger because the oil flowing through it may be used to cool and / or lubricate a generator of the engine 10 configured to provide aircraft power (e.g., an integrated drive generator (IDG) of the engine 10). The depicted fuel management system 1500 is arranged so that fuel reaches the secondary fuel-oil heat exchanger 1004 before the primary fuel-oil heat exchanger 1006.

[0249] The primary fuel-to-oil heat exchanger 1006 and the secondary fuel-to-oil heat exchanger 1004 are configured to carry an oil flow through each in addition to the fuel flow therethrough. The primary fuel-to-oil heat exchanger 1006 and the secondary fuel-to-oil heat exchanger 1004 are configured to allow heat to be transferred between the oil and fuel flowing through them. Under cruise conditions, the average temperature of the oil flow entering the primary fuel-to-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-to-oil heat exchanger 1006, and the average temperature of the oil flow entering the secondary fuel-to-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-to-oil heat exchanger 1004.In this way, 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 stream to an air stream flowing therethrough under flight conditions. The fuel system 1000 includes an electronic controller configured to control the operation of the heat exchangers 1004, 1006.

[0250] The primary fuel-to-oil heat exchanger 1006 may have oil flowing through it that is used to lubricate and / or cool a main gearbox 30 of the gas turbine engine 10 and may therefore be described as a main fuel-to-oil heat exchanger. The secondary fuel-to-oil heat exchanger 1004 may have oil flowing through it that is used to lubricate and / or cool one or more components of a generator of the engine 10 and may therefore be described as a generator fuel-to-oil heat exchanger.

[0251] The two oil streams may be physically separate from each other and optionally also be chemically distinct and / or have a different flow rate. Therefore, a different oil may flow through the primary fuel-oil heat exchanger 1006 than through the secondary fuel-oil heat exchanger 1004.

[0252] In general, at least the majority of the fuel flowing through the secondary fuel-oil heat exchanger 1004 also flows through the main fuel-oil heat exchanger 1006, although both or each heat exchanger 1004, 1006 may be provided with a bypass to allow a portion of the fuel not to flow through the respective heat exchanger, for example in the form of a bypass pipe 1005, as in Fig. 5. A valve (not shown) can determine what portion of the fuel flows through the heat exchanger 1004 and what portion flows through the bypass pipe 1005.

[0253] The two heat exchangers 1004, 1006 can therefore be described as being in series with each other with respect to fuel flow. Generally, at least the majority of the fuel flowing through the secondary fuel-oil heat exchanger 1004 also flows through the main fuel-oil heat exchanger 1006, although each heat exchanger 1004, 1006 may be provided with a bypass to allow a portion of the fuel to not flow through the respective heat exchanger, for example in the form of a bypass pipe 1005. A valve may determine which portion of the fuel flows through the heat exchanger 1004 and which portion flows 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.

[0254] The secondary fuel-oil heat exchanger 1004 and the main fuel-oil heat exchanger 1006 are configured to also convey an oil flow therethrough. The oil flowing through one fuel-oil heat exchanger in the described embodiment is different from the oil flowing through the other fuel-oil heat exchanger, although it is understood that in other embodiments, the same oil may flow through one fuel-oil heat exchanger and then through another fuel-oil heat exchanger. In such implementations, one or more bypass tubes for oil or fuel may again be provided to allow the fluid(s) to bypass one or more heat exchangers.

[0255] The two heat exchangers 1004, 1006 are therefore in the described implementation in separate closed-loop systems with respect to the oil flow, ie the oils flowing through the primary and secondary fuel-oil heat exchangers can be fluidically separated and can be chemically different from each other.

[0256] The fuel system 1000 further includes a modulator valve 1010 located downstream of the primary fuel-to-oil heat exchanger 1006 and arranged to direct at least a portion of the fuel that has exited the primary fuel-to-oil heat exchanger back to the fuel tank 50 via a fuel return line 1011. The modulator valve 1010 is configured to modulate a fuel flow along the fuel return line 1011. The modulator valve 1010 can determine what portion of the fuel is directed back to the fuel tank 50 and what portion continues into the combustion chamber 16. It will be appreciated that the modulator valve 1010 can be located at substantially any suitable position downstream of at least one of the heat exchangers 1004, 1006. For example, the modulator 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 comprise only a single fuel-oil heat exchanger or more than two fuel-oil heat exchangers, and that the modulator valve 1010 may be located downstream of at least one of the heat exchangers (either directly downstream or with one or more intervening components of the fuel system 1000 between the modulator valve 1010 and the heat exchanger).

[0257] The Fig. The heat exchangers shown in Figure 5 and described here are merely examples. The primary and / or secondary heat exchangers can also cool other aircraft systems in addition to those described above.

[0258] In the illustrated example, the modulator valve 1010 is located downstream of the primary fuel-oil heat exchanger 1006. In the illustrated example, the modulator valve 1010 is positioned upstream of the engine fuel pump 1003. The modulator valve 1010 is arranged to allow the return of a controlled amount of fuel to the fuel tank 50.

[0259] The fuel system 1000 includes a plurality of temperature sensors (indicated by the circle symbols surrounding the capital letter T) configured to measure a temperature of the fuel at various locations within the fuel system 1000. In the illustrated example, the fuel system 1000 includes a temperature sensor located within the fuel tank 50 for measuring the fuel temperature within 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 representative of, a fuel temperature within the fuel tank 50.

[0260] 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 modulator valve 1010 and upstream of the combustion chamber 16. Each of the two temperature sensors can be used to determine or represent the temperature of the fuel delivered to the combustion chamber 16 and / or the 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 a position of the modulator 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 provided at any suitable location downstream of the heat exchanger(s) to measure (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.

[0261] The return of fuel to the fuel tank 50 provides a mechanism for controlling the fuel flow of the fuel system 1000, for example, to manage heat loads in the engine 10 and / or to control a temperature of the fuel at various locations in the fuel system 1000 (such as in the fuel tank 50 or upon entering the combustion chamber 16).

[0262] The modulator valve 1010 may be arranged to modulate the fuel flow 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 modulator valve 1010.

[0263] In one example, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of a temperature of the fuel in the fuel tank 50, in Kelvin, to a temperature of the fuel, in Kelvin, delivered to the combustion chamber 16 under cruise 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 as a ΔT ratio (or simply ΔT). A temperature of the fuel 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 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 controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0264] The ΔT 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), where 0 °C corresponds to 273.15 K.

[0265] A fuel temperature 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 in the fuel tank 50 from freezing. Alternatively, the upper limit for the fuel temperature in the fuel tank 50 may be 55°C (328.15 K), 5°C (278.15 K), or 0°C (273.15 K). In some other examples, the upper limit for the fuel temperature in the fuel tank 50 may be 100°C (373.15 K).

[0266] The maximum fuel temperature in the fuel tank can be set by the fuel's flashpoint to reduce the generation of fuel vapors that could potentially ignite in the presence of an ignition source. When fuel is not returned to the tank, it may be advantageous from an engine operating perspective for the fuel in the tank to be colder; however, it must not be so cold that it can cause icing of the fuel in the engine's fuel system. A lower fuel tank temperature allows more of the heat generated in the engine to be redirected back into the fuel, which is more efficient.The inventors recognized that when returning fuel to the tank, the system boundaries extend not only to the engine, but also to the engine and the aircraft fuel tank together, so that a cooler tank is not necessarily more advantageous and the fuel in the tank can be kept at a higher temperature.

[0267] For sustainable aviation fuel (i.e., a fuel consisting of 100% SAF or a blend of SAF), a temperature of the fuel delivered to the combustor 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). More preferably, the temperature of the fuel delivered to the combustor 16 under cruise conditions is typically at least 140°C (413.15 K), or typically between 140°C (413.15 K) and 180°C (453.15 K) or higher (e.g., 200°C or 473.15 K).

[0268] Alternatively, the temperature of the fuel supplied to the combustion chamber 16 under cruise 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 have a fuel temperature that is between any two of these values.

[0269] The temperature of the “fuel delivered to the combustion chamber”, the temperature of the “fuel provided to the combustion chamber” and the temperature of the “fuel supplied to the combustion chamber” are used synonymously here and refer to the same temperature.

[0270] A temperature of the fuel delivered to the combustion chamber 16 under cruise conditions may be defined as an average over at least 5 minutes and optionally over 10 minutes under stable cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in the temperature of the fuel during operation, often increasing the temperature. Each fluctuation must not last more than 5 minutes. The electronic controller of the fuel system 1000 is configured to control the operation of the heat exchangers 1004, 1006 to regulate the temperature of the fuel delivered to the combustion chamber 16.

[0271] The modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is less than 0.56. This value corresponds to less than -54°C / 120°C. The inventors have found that this range can be achieved by using SAFs in which the fuel temperature upon entering the combustion chamber is not limited to 120°C. This might otherwise be the case with fossil kerosene fuel due to the risk of thermal decomposition of the fuel. The minimum temperature of the fuel in the fuel tank is limited to -54°C in this example to avoid the risk of freezing. The value of ΔT in this example (and other similar examples where only an upper limit is specified) would be greater than zero.

[0272] In general, increasing the fuel temperature upon delivery to the combustion chamber can improve the engine's thermodynamic efficiency. The inventors have determined that the use of SAF allows for a higher fuel temperature upon entry into the combustion chamber without the risk of significant thermal degradation of the fuel, which can otherwise occur when using fossil kerosene. The inventors have determined how the engine's fuel / thermal management system can be advantageously controlled to exploit this difference between SAF and kerosene. Increasing the fuel temperature can also be beneficial, as it allows more heat to be transferred from the oil in the heat exchanger, thus improving cooling.

[0273] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is less than 0.53. This value corresponds to less than -54°C / 140°C. This corresponds to a minimum temperature of the fuel supplied to the combustion chamber of 140°C, and the minimum temperature of the fuel in the fuel tank is -54°C. The inventors have determined that operation in this range may be possible with SAF, as it may be advantageous to heat it to a higher temperature to improve thermodynamic efficiency without the risk of thermal breakdown.

[0274] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.48 to 0.56. This corresponds to a value between -54°C / 180°C and -54°C / 120°C. This corresponds to a temperature of the fuel entering the combustion chamber between 120°C and 180°C, while the minimum temperature of the fuel in the fuel tank is -54°C. The maximum limit of 180°C may be advantageous to reduce the risk of fuel decomposition, which can occur even when using SAF.

[0275] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.48 to 0.53. This corresponds to a value between -54°C / 180°C and -54°C / 140°C. This corresponds to a temperature of the fuel entering the combustion chamber between 140°C and 180°C, while the minimum temperature of the fuel in the fuel tank is -54°C. This better utilizes the thermal properties of the SAF to improve thermodynamic efficiency without the risk of fuel decomposition at excessively high temperatures.

[0276] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.48 to 1.00. This corresponds to a temperature range between -54°C / 180°C and 65°C / 65°C. The lower limit corresponds to a fuel temperature upon entering the combustion chamber, which, as explained above, is at the maximum limit of 180°C, while the minimum temperature of the fuel in the tank is limited to -54°C to reduce the risk of freezing. The maximum limit is 1.00 because the heat exchanger(s) in the present examples are not configured to operate in a state where they remove heat from the fuel, i.e., the fuel is not colder upon entering the combustion chamber compared to its temperature in the fuel tank.

[0277] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.48 to 0.90. This corresponds to a temperature range between -54°C / 180°C and 100°C / 140°C. This lower limit of 0.90, compared to the example of the previous paragraph, corresponds to a maximum fuel temperature in the fuel tank of 100°C, as explained above, with the minimum fuel temperature entering the combustion chamber being 140°C to take advantage of the improved thermal properties of the SAF.

[0278] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.48 to 0.82. This corresponds to a temperature range 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 explained above, with the minimum fuel temperature entering the combustion chamber being 140°C, in order to take advantage of the better thermal properties of the SAF and improve thermodynamic efficiency.

[0279] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.56 to 1.00. This corresponds to a temperature range 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 of 140°C or higher upon entering the combustion chamber to utilize the thermal properties of the SAF and improve thermodynamic efficiency, while the minimum temperature at the fuel tank is -54°C to reduce the risk of freezing.

[0280] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.56 to 0.82. This corresponds to a temperature range between -54°C / 140°C and 65°C / 140°C. This corresponds to the same lower limit as 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 explained above, with the minimum fuel temperature entering the combustion chamber being 140°C, in order to take advantage of the better thermal properties of the SAF and improve thermodynamic efficiency.

[0281] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that ΔT is in the range of 0.56 to 0.75. This corresponds to a temperature range between -54°C / 140°C and 65°C / 180°C. This corresponds to the same lower limit as 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 explained above, and the maximum fuel temperature value of 180°C upon entering the combustion chamber, as explained above. The maximum value upon entering 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 them.

[0282] Fig. 8 shows a method 2000 for operating a gas turbine engine 10. The method 2000 includes modulating 2001 the fuel flow along the fuel return line 1011 using the modulator valve 1010 such that a ratio of a fuel temperature in the fuel tank to a temperature of the fuel delivered to the combustion chamber (ΔT) is as in any of the examples described above with respect to the fuel system 1000 or as defined elsewhere herein.

[0283] In another example, the modulator valve 1010 is additionally or alternatively arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up of the engine 10 is less than 0.56. The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 accordingly to achieve this or any other values ​​defined herein from ΔT at cruise to ΔT at start-up.

[0284] Operating the modulator valve in such a way that the ratio of ΔT at cruise to ΔT at start-up of engine 10 is less than 0.56 corresponds to less than a value of ((-54 °C / 120 °C) / 1.00). It therefore corresponds to values ​​that can be achieved by fuel temperatures entering the combustion chamber of 120 °C or higher, which can be achieved by exploiting the improved thermal properties of the SAF compared to fossil fuels. The value of 0.56 corresponds to a minimum value of ΔT, set at at least -54 °C for the fuel in the fuel tank, in order to reduce the risk of freezing when the fuel temperature entering the combustion chamber is at least 120 °C. The maximum value of ΔT at start-up is 1, which corresponds to the maximum that can be achieved without the fuel being hotter at the fuel tank than at the combustion chamber.

[0285] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up 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 the fuel temperature entering the combustion chamber is at least 140°C to further utilize the thermal properties of the SAF to improve thermodynamic efficiency.

[0286] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up of the engine 10 is between 0.48 and 0.56. This corresponds to a value between (-54°C / 180°C) / 1.00) and (-54°C / 120°C) / 1.00). The upper limit of 0.56 corresponds to that explained above. The lower limit of 0.48 corresponds to a maximum fuel temperature upon entering the combustion chamber of 180°C, as explained above, to reduce the risk of thermal decomposition of the fuel, with the minimum fuel temperature in the fuel tank being -54°C to reduce the risk of freezing.

[0287] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up of the engine 10 is between 0.48 and 0.53. This corresponds to a value between (-54°C / 180°C) / 1.00) and (-(54°C / 140°C) / 1.00). The upper limit of 0.53 corresponds to the limit explained above, which further exploits the thermal properties of the SAF. The lower limit of 0.48 corresponds to that explained in the previous paragraph.

[0288] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of cruise ΔT to start-up ΔT 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 cruise ΔT to start-up ΔT ratio can be determined by the following expression: ((min_tank_cruise_flight / max_combustion_chamber_cruise_flight) / (max_tank_start / min_cruise_flight_start@max_tank_cruise_flight))to ((max_tank_cruise_flight / min_combustion_chamber_cruise_flight) / (min_tank_start_flight / max_combustion_chamber_start_flight))

[0289] The lower limit of 0.48 corresponds to the minimum fuel temperature in the tank of -54°C to reduce the risk of freezing, while the maximum fuel temperature entering the combustion chamber at cruise is 180°C to reduce the risk of thermal decomposition of the fuel and improve thermodynamic efficiency. The maximum value of ΔT at takeoff is 1 because the heat exchanger(s) is / are not arranged to cool the fuel before it reaches the combustion chamber. The value of ΔT at takeoff is therefore limited by the minimum fuel temperature entering the combustion chamber, which must not be lower than the corresponding maximum value of the fuel in the tank at which the maximum ΔT 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 entering the combustion chamber 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 entering the combustion chamber during start-up of 160 °C.

[0290] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up 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)). These limits are the same as those in the previous paragraph, except that the maximum temperature of the fuel in the fuel tank is 65°C.

[0291] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up 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, ΔT at take-off may be ΔT at take-off while the aircraft is on the ground. The limits in this paragraph are the same as the previous paragraph, except that the maximum temperature of the fuel in the tank is 100 °C and the maximum temperature of the fuel entering the combustion chamber during take-off is 120 °C.

[0292] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up of the engine 10 is between 0.48 and 1.55. This corresponds to between ((-54°C / 180°C) / (65°C / 65°C)) and ((65°C / 120°C) / (-54°C / 120°C)). In this example, ΔT at take-off may be ΔT at take-off while the aircraft is on the ground. These limits are the same as those in the previous paragraph, except that the maximum temperature of the fuel in the tank is 65°C.

[0293] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up 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, ΔT at take-off may be ΔT at take-off while the aircraft is airborne. These limits are the same as those in the previous paragraph, except that the maximum temperature of the fuel in the tank is 100°C and the maximum temperature of the fuel entering the combustion chamber during take-off is 160°C.

[0294] In some examples, the modulator valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that a ratio of ΔT at cruise to ΔT at start-up 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, ΔT at take-off may be ΔT at take-off while the aircraft is airborne. These limits are the same as those in the previous paragraph, except that the maximum temperature of the fuel in the tank is 65°C.

[0295] In each example described here, the fuel temperature entering the combustion chamber at cruise can be between 120°C and 180°C, and preferably between 140°C and 180°C. In each example described here, the fuel temperature entering the combustion chamber at takeoff can be between -40°C and 160°C. The fuel temperature entering the combustion chamber at ground takeoff can be between -40°C and 120°C. The fuel temperature entering the combustion chamber at airborne takeoff can be between 0°C and 160°C.

[0296] The temperature range in the previous paragraph may correspond to a ground takeoff, where the engine starts and all components are at ambient temperature. For an airborne (i.e., in-flight) takeoff, the temperature range in the previous paragraph corresponds to a reignition scenario. In this case, the engine was previously running and is then restarted. This may include the aircraft descending from cruising altitude to a windmill reignition altitude. During this period, the temperature of the engine's core components decreases, but some residual heat may still be present in the engine at the time of reignition. The temperature ranges for a ground restart and an air restart are therefore different.

[0297] Fig. 9 shows a method 2100 for operating a gas turbine engine 10. The method 2100 includes modulating 2101 the fuel flow along the fuel return line 1011 using the modulator valve 1010 such that the ratio of ΔT at cruise to ΔT at start-up of the engine 10 is as defined in any of the examples above or elsewhere herein.

[0298] In another example, the modulator valve 1010 is additionally or alternatively arranged to initiate the return of fuel to the fuel tank 50 when the fuel delivered to the combustion chamber 16 (or having flowed through one or both heat exchangers 1004, 1006) has a temperature of at least 120°C. At fuel temperatures below 120°C, the modulator valve 1010 may prevent the return of fuel to the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 accordingly. In other examples, the modulator valve may be arranged to initiate the return of fuel to the fuel tank 50 when the fuel delivered to the combustion chamber 16 (or having flowed through one or both heat exchangers 1004, 1006) has a temperature of at least 140°C.This allows the thermal properties of the SAF to be further exploited. In yet other examples, the modulator valve may be arranged to initiate the return of fuel to the fuel tank when the fuel that has flowed through one or both heat exchangers has a temperature between 120°C and 180°C, and preferably has a temperature between 140°C and 180°C.

[0299] The temperature of the fuel is measured downstream of one or both heat exchangers 1004, 1006. In this example, the temperature of the fuel is measured using a temperature sensor in the fuel return line 1011, although it is understood that the temperature of the fuel may be measured using a temperature sensor located virtually anywhere downstream of one or both heat exchangers 1004, 1006. The temperature of the fuel that has flowed through one or both heat exchangers 1004, 1006 may be used to determine, or be representative of, the temperature of the fuel returned to the fuel tank 50.The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 based at least in part on temperature data from the temperature sensors located downstream of one or both of the heat exchangers 1004, 1006.

[0300] In this example, the modulator valve 1010 is arranged to initiate the return of fuel to the fuel tank 50 when the fuel that has flowed through one or both of the heat exchangers 1004, 1006 has 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 required.

[0301] 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 degree of coking of the fuel, an oxygen content of the fuel, and a sulfur content of the fuel.

[0302] The operating conditions may be detected onboard, for example during flight, using one or more sensors and / or other measurements. The data from the sensor(s) may be transmitted to the electronic control unit of the fuel system 1000 to control the operation of the modulator valve 1010. Alternatively, the operating conditions may be interrogated, manually entered, or transmitted to the aircraft 1 (for example, to the electronic engine control unit of the fuel system 1000) before or during the flight.

[0303] For example, trace substances or species, either naturally present in the fuel or added as tracers, can be used to determine fuel properties, such as the proportion 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 can be used as a basis for determining various fuel properties, including the aromatic content of the fuel, the oxygen content of the fuel, the thermal stability of the fuel, and the degree of coking of the fuel, for example, by measuring the formation of surface deposits on the piezoelectric crystal that result in a change in the vibration mode.Other fuel properties, such as sulfur content, naphthalene content, aromatic hydrocarbon content, and hydrogen-carbon ratio, can be determined by measuring the substances present in the exhaust gases emitted by the gas turbine 10 during use (e.g., in flight).

[0304] In this example, the modulator valve 1010 is also arranged to modulate the fuel flow along the fuel return line 1011 based at least in part on the amount of fuel remaining in the fuel tank, although this is not required.

[0305] The modulator valve 1010 is additionally or alternatively arranged to modulate the fuel flow along the fuel return line 1011 such that an equilibrium temperature of the fuel in the fuel tank 50 after the fuel has been returned to the fuel tank 50 does not exceed a maximum allowable temperature in the fuel tank 50 (e.g., 65°C or 100°C). The electronic controller of the fuel system 1000 can determine how long the modulator valve 1010 can remain open and / or how much fuel can 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 modulator valve 1010). An example of logic that can be used to operate the modulator valve 1010 in this manner is provided below: (Mtank+MFRTT)⋅Cp(Teq)⋅Teq=Mtank⋅Cp(Ttank)⋅Ttank+MFRTT⋅Cp(TFRTT)⋅TFRTT where M tank a fuel mass in the fuel tank is 50, M FRTT is a fuel mass that is returned to the fuel tank 50 via the fuel return line 1011, T eq an equilibrium temperature of the mixed fuel in the fuel tank 50 after the fuel is returned to the fuel tank 50, T tank a current temperature of the fuel in the fuel tank 50 before mixing is T FRTT is the temperature of the fuel returned to the fuel tank 50 via the fuel return line 1011, and C p is the specific heat capacity of the fuel at the corresponding temperature (for example T eq , T tank , T FRTT). 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 being 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 modulator valve 1010 may be operated accordingly to return the required mass of fuel to the fuel tank 50.

[0306] The temperature of the fuel that has flowed through one or both of heat exchangers 1004, 1006 may be used to determine, or be representative of, the temperature of the fuel returned to fuel tank 50. In this example, the temperature of the fuel returned to fuel tank 50 is measured using a temperature sensor in fuel return line 1011, although it is understood that the temperature of the fuel returned to fuel tank 50 may be measured using a temperature sensor located virtually anywhere downstream of one or both of heat exchangers 1004, 1006. The temperature of the fuel in fuel tank 50 may be measured using a temperature sensor located in fuel tank 50 or a temperature sensor located downstream of fuel tank 50 and upstream of heat exchangers 1004, 1006.The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0307] It will be appreciated that any suitable alternative logic may be used to control the operation of the modulator valve 1010 (e.g., based at least in part on the amount of fuel remaining in the fuel tank 50).

[0308] Fig. 10 shows a method 2200 for operating a gas turbine engine 10. The method 2200 includes initiating 2201 the return of fuel to the fuel tank using the modulator valve 1010 when the fuel flowing through the heat exchanger(s) has a temperature of at least 120°C or is within one of the other ranges as described above with respect to the fuel system 1000 or as described elsewhere herein.

[0309] In another example, the modulator valve 1010 is additionally or alternatively arranged to modulate the fuel flow along the fuel return line 1011 such that a ratio of the fuel mass returned to the tank 50 to the fuel mass delivered to the combustion chamber 16 is between 0 and 9 under cruise conditions. The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010. The modulator valve 1010 may be arranged to modulate the fuel flow along the fuel return line 1011 such that the ratio of the fuel mass returned to the tank 50 to the fuel mass delivered to the combustion chamber 16 is greater than 4 and less than or equal to 9 under cruise conditions, and preferably greater than 4 and less than or equal to 4.9 under cruise conditions.

[0310] SAF typically has a higher calorific value than conventional fuels such as kerosene or kerosene-based fuels. For example, SAF can 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. Therefore, a lower mass flow of SAF is required to provide the same amount of fuel power input to the combustion chamber as when using conventional fuels.

[0311] Furthermore, SAF typically exhibits higher thermal stability than conventional fuels and can therefore operate at a higher temperature before fuel degradation (due to thermal decomposition) occurs. SAF can therefore absorb more heat than conventional fuels before fuel degradation occurs.

[0312] The return of fuel to the fuel tank 50 typically occurs when additional fuel is used for purposes other than combustion, such as to manage the heat load in the engine 10. For example, fuel may be used to cool oil in the engine 10 via the heat exchangers 1004, 1006 and / or to drive one or more actuators to operate other components in the engine 10 or the aircraft 1 in general. Because SAF can absorb more heat than conventional fuels and has a higher calorific value, the inventors have determined that a different ratio between the mass of fuel returned to the tank and the mass of fuel delivered to the combustion system may result in an increase in performance.For example, by using SAF, the higher calorific value of the fuel can result in less fuel mass being burned and thus a larger percentage of fuel being returned to the tank than when using conventional fossil fuels.

[0313] Fig. 11 shows a method 2300 for operating a gas turbine engine 10. The method 2300 includes modulating 2301 the fuel flow along the fuel return line 1011 such that a ratio of the fuel mass returned to the tank 50 to the fuel mass delivered to the combustor 16 is between 0 and 9 under cruise conditions, as described above with respect to the fuel system 1000. The ratio of the fuel mass returned to the tank 50 to the fuel mass delivered to the combustor 16 may be in or within any of the other ranges defined above or elsewhere herein.

[0314] In another example, the modulator valve 1010 is additionally or alternatively arranged to prevent the return of fuel having a temperature of 180°C or higher to the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 accordingly.

[0315] The temperature of the fuel is measured downstream of one or both heat exchangers 1004, 1006. In this example, the temperature of the fuel is measured using a temperature sensor in the fuel return line 1011, although it is apparent that the temperature of the fuel can be measured using a temperature sensor located virtually anywhere downstream of one or both heat exchangers 1004, 1006. The temperature of the fuel that has flowed through one or both heat exchangers 1004, 1006 can be used to determine, or be representative of, the temperature of the fuel in the fuel return line 1011.The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve based at least in part on temperature data from the temperature sensors located downstream of one or both of the heat exchangers 1004, 1006.

[0316] In this example, the modulator valve 1010 is also arranged to prevent the return of fuel to the fuel tank when the fuel temperature in the fuel tank 50 reaches a predetermined upper threshold temperature. As described above, the fuel temperature in the fuel tank 50 is typically maintained between -54°C (219.15 K) and 100°C (373.15 K). More preferably, the temperature should be maintained between -54°C (219.15 K) and 65°C (338.15 K). The upper limit for the temperature of the fuel in the fuel tank 50 may be any suitable value between the temperature pairs in the two sets above; for example, the upper limit may be 55°C (328.15 K), 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.

[0317] By preventing the return of fuel to the fuel tank 50 when the fuel in the fuel tank 50 has reached the predetermined upper threshold temperature, the threshold temperature will not be exceeded as a result of the return of fuel that has flowed 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 controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors that provide data indicative of a temperature of the fuel returned to the fuel tank 50 and a temperature sensor that provides data indicative of a temperature of the fuel in the fuel tank 50.

[0318] In this example, the modulator valve 1010 is also arranged to modulate the fuel flow along the fuel return line 1011 based at least in part on the amount of fuel remaining in the fuel tank, although this is not required.

[0319] The modulator valve 1010 may be arranged to enable the return of fuel to the fuel tank 50 when a temperature of the fuel returned to the fuel tank 50 is below 180°C and a temperature of the fuel in the fuel tank 50 is below the predetermined upper threshold temperature. If both conditions are met, the modulator valve 1010 is arranged to modulate the fuel flow 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 the predetermined upper threshold temperature.The electronic controller of fuel system 1000 can determine how long the modulator valve 1010 can remain open and / or how much fuel can be returned to the fuel tank 50 (and optionally, 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 can be used to operate the modulator valve 1010 in this manner is provided below:. (Mtank+MFRTT)⋅Cp(Teq)⋅Teq=Mtank⋅Cp(Ttank)+MFRTT⋅Cp(TFRTT)⋅TFRTT where M tank a fuel mass in the fuel tank is 50, M FRTT is a fuel mass that is returned to the fuel tank 50 via the fuel return line 1011, T eq an equilibrium temperature of the mixed fuel in the fuel tank 50 after the fuel is returned to the fuel tank 50, T tanka current temperature of the fuel in the fuel tank 50 before mixing is T FRTT is a current temperature of the fuel returned to the fuel tank 50 via the fuel return line 1011, and C p is the specific heat capacity of the fuel at the corresponding temperature (for example T eq , T tank , T FRTT). 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 being returned to the fuel tank 50 along the fuel return line 1011, ensuring that the equilibrium temperature does not exceed the 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 modulator valve 1010 may be operated accordingly to return the required mass of fuel to the fuel tank 50.

[0320] The temperature of the fuel that has flowed through one or both of heat exchangers 1004, 1006 may be used to determine, or be representative of, the temperature of the fuel returned to fuel tank 50. In this example, the temperature of the fuel returned to fuel tank 50 is measured using a temperature sensor in fuel return line 1011, although it is understood that the temperature of the fuel returned to fuel tank 50 may be measured using a temperature sensor located virtually anywhere downstream of one or both of heat exchangers 1004, 1006. The temperature of the fuel in fuel tank 50 may be measured using a temperature sensor located in fuel tank 50 or a temperature sensor located downstream of fuel tank 50 and upstream of heat exchangers 1004, 1006.The electronic controller of the fuel system 1000 is configured to control the operation of the modulator valve 1010 based at least in part on temperature data from at least one of the temperature sensors.

[0321] It will be appreciated that any suitable alternative logic may be used to control the operation of the modulator valve 1010 (e.g., based at least in part on the amount of fuel remaining in the fuel tank 50).

[0322] Fig. 12 shows a method 2400 for operating a gas turbine engine 10. The method 2400 includes preventing 2401 the return of fuel having a temperature of 180°C or higher to the fuel tank 50 using the modulator valve 1010, as described above with respect to the fuel system 1000. The method 2400 may include preventing 2401 the return of fuel having a temperature within any of the ranges defined above or elsewhere herein to the fuel tank 50 using the modulator valve 1010. The method may include preventing the return of the fuel if doing so would exceed the upper tank temperature limit defined above or elsewhere herein.

[0323] 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 described below from Fig. 13 if it is equipped with a suitable fuel recirculation line.

[0324] An exemplary fuel system 1100 including a fuel flow path from the fuel tank 50 to other components of the fuel system 1100 is schematically shown in Fig. 6. The fuel system 1100 is essentially the same as described above with reference to Fig. 5, with like reference numerals designating like elements.

[0325] The fuel system 1100 includes a fuel recirculation line 1111. The fuel recirculation line 1111 is arranged to recirculate 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 main 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 modulator valve 1110 arranged to modulate fuel flow along the fuel recirculation line 1111 between the inlet and 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 required. The fuel system 1100 may also include a fuel return line (not shown for clarity), as described above with respect to the system shown in FIG. Fig. 5, although this is not mandatory.

[0326] In the example shown, the first point, or inlet, of fuel recirculation line 1111 is located downstream of primary fuel-to-oil heat exchanger 1106 in the fuel flow path. The first point is downstream of fuel pump 1103. Modulator valve 1110 is also located in the fuel flow path, although it is understood that modulator valve 1110 could instead be located along fuel recirculation line 1111. The second point, or outlet, of fuel recirculation line 1111 is upstream of primary fuel-to-oil heat exchanger 1106.

[0327] Alternatively, the inlet and outlet of fuel recirculation line 1111 may be located at any suitable location in the fuel flow path. For example, the inlet of fuel recirculation line 1111 may be located upstream of fuel pump 1103, or downstream of secondary fuel-oil heat exchanger 1104, or downstream of fuel pump 1102. The outlet of fuel recirculation line 1111 may be located upstream of fuel pump 1103, or upstream of secondary fuel-oil heat exchanger 1104, or upstream of fuel pump 1102. If the outlet of fuel recirculation line 1111 is upstream of fuel pump 1102, the inlet of the recirculation line may be upstream of pump 1103.It will also be appreciated that the fuel system 1100 may alternatively comprise only a single fuel-to-oil heat exchanger or more than two fuel-to-oil heat exchangers, and that the first or second point of the fuel recirculation line 1111 may be located at any suitable position (upstream or downstream, either directly or indirectly with one or more intervening components in the main fuel flow path) with respect to the one or more heat exchangers. One fuel recirculation line 1111 may be provided for any one fuel-to-oil heat exchanger.

[0328] The fuel recirculation line 1111 is arranged to allow some fuel to flow multiple times along at least a portion of the fuel flow path (that is, to recirculate some of the fuel along at least a portion of the fuel flow path). In the example shown, the modulator valve 1110 is arranged to modulate the fuel flow along the fuel recirculation line 1111 such that an amount of fuel flows through the primary fuel-oil heat exchanger 1106 multiple times before being delivered to the combustion chamber 16. This may potentially transfer additional heat to the same fuel and increase the temperature of the fuel before the fuel is delivered to the combustion chamber 16. The temperature of the fuel may be increased to at least 120°C before the fuel is delivered to the combustion chamber 16.Recirculating the fuel through the fuel recirculation line 1111 using the modulator valve 1110 may therefore enable the delivery of hotter fuel to the combustion chamber 16, which may increase the efficiency of the engine 10. Furthermore, an amount (e.g., a mass) of fuel used to manage heat loads within the engine 10 (e.g., via the heat exchangers 1104, 1106) may be reduced, which in turn may reduce (or, depending on operating conditions, eliminate) an amount of fuel that must be returned to the fuel tank 50 to manage heat loads within the engine 10. By reducing or avoiding an amount of fuel returned to the fuel tank 50, it can also be achieved that the fuel in the fuel tank 50 is cooler before it enters the fuel flow path, which can improve the cooling of the oil in the fuel-oil heat exchangers 1104, 1106.

[0329] The modulator valve 1110 and / or the inlet of the fuel recirculation tube 1111 may be located downstream of the fuel pump 1103 so that the recirculated fuel flows through both the primary fuel-oil heat exchanger 1106 and the fuel pump 1103 multiple times, rather than just the primary heat exchanger 1106. By recirculating fuel through the fuel pump 1103, more adjustable control of the fuel flow rate to the combustion chamber 16 may be enabled for a given shaft speed of the engine 10, noting that the pump speed (or a limited set of pump speed options) is often set by the shaft speed (e.g., the pump speed may be proportional to the shaft speed).For example, if the fuel flow rate required to maintain current aircraft thrust is lower than the pump speed for the required shaft speed, a larger portion of fuel can be recirculated through the pump so that the flow rate through the pump is higher than the flow rate to the combustor 16. Therefore, a lower flow rate of fuel to the combustor 16 for a higher calorific value fuel can be provided without changing the pump speed by instead increasing the amount of recirculation. Typically, no fuel is wasted, as it is always recirculated and burned in the combustor.

[0330] In the example shown, the modulator valve 1110 is arranged to actively control an amount of fuel recirculated along the fuel recirculation line 1111. The electronic control device of the fuel system 1000 is configured to control the operation of the modulator valve 1110. The active control may be performed based on one or more parameters, such as: • Core shaft speed and engine thrust requirements; • 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; • Fuel calorific value; • Fan speed; • Fuel flow rate to the combustion chamber (commonly referred to as WFE - Weight of Main Engine Fuel Flow); • Fuel viscosity; and • Main / engine fuel pump speed or speed options.

[0331] Additionally or alternatively, a speed of the fuel pump 1103 may be adjusted to either accelerate the fuel flow rate (and thus reduce the heat transfer per unit volume of fuel flowing through the heat exchangers 1104, 1106) or reduce the fuel flow rate (and thus increase the heat transfer per unit volume of fuel flowing through the heat exchangers 1104, 1106).

[0332] Alternatively, control of the amount of fuel recirculated along the fuel recirculation line 1111 using the modulator valve 1110 may not be active. Rather, a fixed or unchanging portion of the fuel in the main fuel flow path may be recirculated along the fuel recirculation line 1111 via the modulator valve 1110; for example, the ratio of the recirculated fuel mass to the fuel mass delivered to the combustion chamber may be between 0 and 9. More preferably, the ratio between the recirculated fuel mass and the fuel mass delivered to the combustion chamber may be greater than 4 and less than or equal to 9 under cruise conditions, and even more preferably, the ratio may be greater than 4 and less than or equal to 4.9 under cruise conditions.It may therefore be the same ratio as the ratio defined above for returning the fuel to the tank, so that everything described herein that applies to returning the fuel to the tank may also apply to the present example, in which the fuel is instead recirculated via the fuel recirculation line 1111.

[0333] In some examples, fuel recirculation line 1111 is arranged to supply fuel to one or more additional aircraft and / or engine mechanisms (not shown) before re-entering the fuel flow path. The additional aircraft mechanism(s) may therefore be located on fuel recirculation line 1111. The additional aircraft mechanisms may include an engine nacelle anti-icing system, actuators, a turbine casing cooling system, or any other suitable aircraft mechanism.

[0334] Fig. 13 shows a method 2500 for operating a gas turbine engine 10. The method 2500 includes modulating 2501 the fuel flow along the fuel recirculation line 1111 as described above with respect to the fuel system 1100.

[0335] An exemplary fuel system 1200 including a fuel flow path from the fuel tank 50 to other components of the fuel system 1200 is schematically illustrated in Fig. 7. In the example shown, the engine 10 is a direct drive engine, as in Fig. 4. The fuel system 1200 from Fig. 7 can, however, also be used with other engine architectures, for example, a gearbox architecture. The fuel system 1200 includes both the fuel supply system (comprising the fuel tank 50 and the pump 1202), which supplies fuel to the engine, 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 can manage the fuel supply to multiple engines and can, for example, be a duplicate of various ones of the Fig. 6 for the other engine(s).

[0336] A low-pressure fuel supply pump 1202 pumps fuel from the fuel tank 50 to the gas turbine engine 10. 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 may be described as a main or high-pressure fuel pump. At least a portion of the fuel then flows through a secondary fuel-to-oil heat exchanger 1206 via a conduit 1207 branching from the main fuel flow path (between the fuel tank 50 and the combustion chamber 16), and at least a portion of the fuel flows toward the combustion chamber 16 without flowing through the secondary fuel-to-oil heat exchanger 1206. The primary fuel-oil heat exchanger 1204 may be described as a main fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1206 may be described as a servo fuel-oil heat exchanger.The fuel management system 1600 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1204 before the secondary fuel-oil heat exchanger 1206.

[0337] In the Fig. 7, the portion of the fuel that has flowed through the secondary fuel-oil heat exchanger 1206 then flows to the combustion chamber 16 of the gas turbine engine 10, where it rejoins the fuel that has flowed only through the primary fuel-oil heat exchanger 1204.

[0338] The primary fuel-to-oil heat exchanger 1204 and the secondary fuel-to-oil heat exchanger 1206 are configured to also convey an oil flow therethrough. The primary fuel-to-oil heat exchanger 1204 and the secondary fuel-to-oil heat exchanger 1206 are configured to allow heat to be transferred between the oil and fuel flowing through them. Under cruise conditions, the average temperature of the oil flow entering the primary fuel-to-oil heat exchanger 1204 and the secondary fuel-to-oil heat exchanger 1206 is higher than the average temperature of the fuel entering the primary fuel-to-oil heat exchanger 1204 and the secondary fuel-to-oil heat exchanger 1206, respectively.In this way, the primary fuel-oil heat exchanger 1204 and the secondary fuel-oil heat exchanger 1206 are configured to transfer thermal energy from an oil stream to an air stream flowing therethrough under cruise conditions. The fuel system 1200 includes an electronic controller configured to control the operation of the heat exchangers 1204, 1206.

[0339] In various configurations of the fuel management system 1600, the oil flows through the secondary fuel-to-oil heat exchanger 1206 before flowing through the primary fuel-to-oil heat exchanger 1204 and does not flow through engine components that would increase its temperature therebetween. Therefore, the oil is hotter entering the secondary fuel-to-oil heat exchanger 1206 than entering the primary fuel-to-oil heat exchanger 1204. Conversely, the fuel flows through the primary fuel-to-oil heat exchanger 1204 before passing through the secondary fuel-to-oil heat exchanger 1206. In this configuration, the fuel temperature exiting the secondary fuel-to-oil heat exchanger 1206 is higher than the fuel temperature exiting the primary fuel-to-oil heat exchanger 1204.The engine fuel pump 1203 may alternatively be placed upstream of the secondary fuel-oil heat exchanger 1206 (or upstream of where the line 1207 rejoins the main fuel flow path), meaning that it is not exposed to these further elevated fuel temperatures.

[0340] The fuel management system 1600 may be provided with a bypass to prevent a portion of the fuel from flowing through the respective heat exchanger, for example in the form of a bypass pipe having 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 (not shown) can determine what portion of the fuel flows through the heat exchanger 1204 and what portion flows through the bypass pipe 1005.

[0341] The fuel system 1200 further includes a modulator valve 1210 located downstream of the primary fuel-to-oil heat exchanger 1204 and arranged to divert or direct at least a portion of the fuel that has exited the primary fuel-to-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 fuel flow along the fuel return line 1211. The modulator valve 1210 can determine what portion of the fuel is directed back to the fuel tank 50 and what portion continues into the combustion chamber 16. It will be appreciated that the modulator valve 1210 can be located at substantially any suitable 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 conduit 1207 rejoins the main fuel flow path). It will also be appreciated that the fuel system 1200 may alternatively include only a single fuel-to-oil heat exchanger or more than two fuel-to-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 intervening components of the fuel system 1200 between the modulator valve 1210 and the heat exchanger). The primary and secondary heat exchangers described above are merely an example and may provide cooling for any suitable engine systems or components.

[0342] In the illustrated example, the modulator valve 1210 is located downstream of the primary fuel-oil heat exchanger 1204, the main fuel pump 1203, the inlet of line 1207 leading to the secondary fuel-oil heat exchanger 1206, and the outlet of line 1207 leading from the secondary fuel-oil heat exchanger 1206 back into the main fuel flow path. It will be appreciated that the modulator valve 1210 may alternatively be located upstream of the fuel pump 1203.

[0343] The fuel system 1200 includes a plurality of temperature sensors (indicated by the circle symbols surrounding the capital letter T) configured to measure a temperature of the fuel at various locations within the fuel system 1200. In the illustrated example, the fuel system 1200 includes a temperature sensor located within the fuel tank 50 for measuring the fuel temperature within 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 1204, 1206, which may be used to determine, or representative of, a fuel temperature within the fuel tank 50.

[0344] In the illustrated example, fuel system 1200 includes two temperature sensors located downstream of heat exchangers 1204, 1206. One of the temperature sensors is located in fuel return line 1211, while the other temperature sensor is located downstream of modulator valve 1210 and upstream of combustion chamber 16. Each of the two temperature sensors can be used to determine or represent the temperature of the fuel delivered to combustion chamber 16 and / or the temperature of the fuel returned to fuel tank 50. It will be appreciated that only one of the temperature sensors located downstream of heat exchangers 1204, 1206 may be provided.Depending on a position of the modulator 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 provided at any suitable location downstream of the heat exchanger(s) to measure (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.

[0345] The return of fuel to the fuel tank 50 provides a mechanism for controlling the fuel flow of the fuel system 1200, for example, to manage heat loads in the engine 10 and / or to control a temperature of the fuel at various locations in the fuel system 1200 (such as in the fuel tank 50 or upon entering the combustion chamber 16).

[0346] The modulator valve 1210 may be arranged to modulate the fuel flow along the fuel return line 1211 to the fuel tank 50 in any suitable manner, for example as described above with respect to the Fig. 5. The fuel system 1200 includes an electronic control device configured to control the operation (e.g., opening and closing) of the modulator valve 1210. It will also be appreciated that one or more of the Fig. 8 to 12 and described above with respect to the fuel system 1000, using the method 2000, 2100, 2200, 2300, 2400 shown in Fig. 7. The fuel system 1200 may also perform the method 2500 of Fig. 13 if it is equipped with a suitable fuel recirculation line as described below.

[0347] An exemplary fuel system 1300 including a fuel flow path from the fuel tank 50 to other components of the fuel system 1300 is schematically illustrated in Fig. 14. The fuel system 1300 is essentially the same as described above with reference to Fig. 7, with like reference numerals designating like elements.

[0348] The fuel system 1300 includes a fuel recirculation line 1311. The fuel recirculation line 1311 is arranged to recirculate 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 running 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 modulator valve 1310 arranged to modulate fuel flow along the fuel recirculation line 1311 between the inlet and 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 required. The fuel system 1300 may also include a fuel return line (not shown for clarity), as described above with respect to the system shown in FIG. Fig. 7, although this is not mandatory.

[0349] In the example shown, the first point or inlet of the fuel recirculation line 1311 is located downstream of the primary fuel-to-oil heat exchanger 1304 in the fuel flow path. The first point is downstream of the fuel pump 1303. The first point is downstream of the inlet of line 1307 leading to the secondary fuel-to-oil heat exchanger 1306 and downstream of the outlet of line 1307 leading from the secondary fuel-to-oil heat exchanger 1306 back into the main fuel flow path (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 upstream of the flow paths to and from the secondary heat exchanger.The modulator valve 1310 is also located on the main fuel flow path, although it should be appreciated that the modulator valve 1310 may instead be located elsewhere in 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.

[0350] The inlet and outlet of fuel recirculation line 1311 may alternatively be located at any suitable location in the fuel flow path. For example, the inlet of fuel recirculation line 1311 may be located downstream of primary heat exchanger 1304 but upstream of fuel pump 1303, or downstream of primary heat exchanger 1304 and fuel pump 1303 but upstream of the inlet of line 1307 leading to secondary heat exchanger 1306, or downstream of the outlet of line 1307 rejoining the main fuel flow path from 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 only a single fuel-to-oil heat exchanger or more than two fuel-to-oil heat exchangers, and that the first or second point of the fuel recirculation line 1311 may be located at any suitable position (upstream or downstream, either directly or indirectly with one or more intervening components in the fuel flow path) with respect to the one or more heat exchangers. One fuel recirculation line 1311 may be provided for each fuel-to-oil heat exchanger.

[0351] The fuel recirculation line 1311 is arranged so that a portion of the fuel can flow multiple times along at least a portion of the fuel flow path (that is, a portion of the fuel is recirculated along at least a portion of the fuel flow path) before being delivered to the combustion chamber 16. In the example shown, the modulator valve 1310 is arranged to modulate the fuel flow along the fuel recirculation line 1311 so that an amount of fuel flows through the primary fuel-oil heat exchanger 1304 multiple times before being delivered to the combustion chamber 16. This may potentially transfer additional heat to the same fuel and increase the temperature of the fuel before the fuel is delivered to the combustion chamber 16. The temperature of the fuel may be increased to at least 120°C before the fuel is delivered to the combustion chamber 16.In some examples, the temperature of the fuel may be increased to at least 140°C before the fuel is delivered to the combustion chamber 16, or to a value in 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 modulator valve 1310 may therefore allow for the delivery of hotter fuel to the combustion chamber 16, which may increase the efficiency of the engine 10. Furthermore, an amount (e.g., a mass) of fuel used to manage heat loads within the engine 10 (e.g., via the heat exchangers 1304, 1306) may be reduced, which in turn may reduce (or, depending on operating conditions, eliminate) an amount of fuel that must be returned to the fuel tank 50 to manage heat loads within the engine 10.By reducing or avoiding an amount of fuel returned to the fuel tank 50, it can also be achieved that the fuel in the fuel tank 50 is cooler before it enters the fuel flow path, which can improve the cooling of the oil in the fuel-oil heat exchangers 1304, 1306.

[0352] In the example shown, the modulator valve 1310 and / or the inlet of the fuel recirculation tube 1311 is also located downstream of the fuel pump 1303, such that the recirculated fuel flows through both the primary fuel-oil heat exchanger 1304 and the fuel pump 1303 multiple times. By recirculating fuel through the fuel pump 1303, more adjustable control of the fuel flow rate to the combustion chamber 16 may be enabled for a given shaft speed of the engine 10, noting that the pump speed (or a limited set of pump speed options) is often set by the shaft speed (e.g., the pump speed may be proportional to the shaft speed).For example, if the fuel flow rate required to maintain current aircraft thrust is lower than the pump speed for the required shaft speed, a larger proportion of fuel may be recirculated through the pump so that the flow rate through the pump is higher than the flow rate to the combustor 16. Therefore, a lower flow rate of fuel to the combustor 16 for a higher calorific value fuel may be provided without changing the pump speed by instead increasing the amount of recirculation.

[0353] The modulator valve 1310 may be arranged to modulate the fuel flow along the fuel recirculation line 1311 in any suitable manner, for example as described above with respect to the Fig. 7. The fuel system 1300 includes an electronic control device configured to control the operation (e.g., opening and closing) of the modulator valve 1310. It is also apparent that the Fig. 13 and described above with respect to the fuel system 1100, using the method 2500 shown in Fig. 14 shown fuel system 1300.

[0354] In some examples, fuel recirculation line 1311 is arranged to supply fuel to one or more additional engines and / or aircraft mechanisms (not shown) before rejoining the fuel flow path. The additional aircraft mechanism(s) may therefore be located on fuel recirculation line 1311. The additional aircraft mechanisms may include an engine nacelle anti-icing system, actuators, a turbine casing cooling system, or any other suitable aircraft mechanism.

[0355] Further exemplary fuel systems 1400, 1700 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 are shown schematically in the Fig. 15 and Fig. 16. The fuel system 1400 is similar to the fuel system described above with reference to Fig. 5 and Fig. 6, wherein like reference numerals indicate like elements. The fuel system 1700 is similar to the fuel systems described above with reference to Fig. 7 and Fig. 14 described fuel systems 1200, 1300, wherein like reference numbers designate like elements.

[0356] In the fuel systems 1400, 1700, a modulator valve 1410, 1710 is arranged to divert or direct at least a portion of the fuel that has left a fuel-oil heat exchanger back to the fuel tank 50 and / or to return fuel to an upstream location in the fuel flow path. Therefore, the method of one or more of the Fig. Perform steps 8 to 13.

[0357] In the Fig. 15, the modulator valve 1410 of the fuel system 1400 is located downstream of the secondary fuel-to-oil heat exchanger 1406 and the pump 1403. The modulator valve 1410 is arranged to divert or redirect 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 arranged to recirculate fuel from the pump 1403 via a fuel return line 1411a to a point upstream of the secondary heat exchanger 1406. It will also be appreciated that the modulator valve 1410 may alternatively be located at any suitable location in the fuel system 1400, for example, along the fuel flow path.

[0358] The modulator valve 1410 may be arranged to modulate the fuel flow 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 Fig. 5, Fig. 6, Fig. 7 and Fig. 14. The fuel system 1400 includes an electronic control device configured to control the operation (e.g., opening and closing) of the modulator valve 1410. It is also apparent that the Fig. 8 to 13 and described above with respect to the fuel systems 1000, 1100, 1200, 1300 in the same way using the method 2000, 2100, 2200, 2300, 2400, 2500 in Fig. 15 shown fuel system 1400.

[0359] In the Fig. 16, the modulator valve 1710 of the fuel system 1700 is located downstream of the secondary fuel-to-oil heat exchanger 1706. Fuel exiting the secondary heat exchanger 1706 does not reenter the main fuel flow path (between the fuel tank 50 and the combustion chamber 16). The modulator valve 1710 is arranged to divert or redirect 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 arranged to recirculate the fuel exiting the secondary heat exchanger 1706 via a fuel return line 1711a to a point upstream of the primary fuel-to-oil heat exchanger 1704 on the main fuel flow path.It will also be appreciated that the modulator valve 1710 may alternatively be located at any suitable location, for example, on the main fuel flow path, so that the fuel exiting the secondary heat exchanger 1706 can re-enter the main fuel flow path before encountering the modulator valve 1710.

[0360] The modulator valve 1710 may be arranged to modulate the fuel flow 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 Fig. 5, Fig. 6, Fig. 7 and Fig. 14. The fuel system 1700 includes an electronic control device configured to control the operation (e.g., opening and closing) of the modulator valve 1710. It is also apparent that the Fig. 8 to 13 and described above with respect to the fuel systems 1000, 1100, 1200, 1300 in the same way using the method 2000, 2100, 2200, 2300, 2400, 2500 in Fig. 16 shown fuel system 1700.

[0361] Fuel temperatures upon entering the combustion chamber under cruise conditions, or upon initiation of fuel return, as defined elsewhere herein, may be defined as an average over at least 5 minutes, 10 minutes, or 30 minutes under steady-state conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in the fuel temperature during operation, often increasing in temperature. Each fluctuation must not last more than 5 minutes.

[0362] Fig. Figure 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 to the gas turbine engines 10 from the fuel tank 50. In this example, the fuel tank 50 consists of a series of interconnected fuel tanks. In the example in Fig.17, the fuel tank consists of a main fuel tank located in the fuselage and a smaller fuel tank located in each wing. In other examples, an aircraft 1 may have only a single fuel tank. It is obvious that many different tank layouts are possible and that the illustrated examples are for ease of description and are not intended as a limitation.

[0363] It should 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 in cases of mutual exclusion, each of the features may be used separately or in combination with any other features, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

[1] 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 flowed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulator valve (1010; 1210; 1410; 1710) arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) such that a ratio of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is between 0 and 9 under cruise conditions. [2] A gas turbine engine (10) according to claim 1, wherein the modulator 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 of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is between 2.3 and 9 under cruise conditions. [3] A gas turbine engine (10) according to claim 1 or claim 2, wherein the modulator 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 of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is greater than 4 and less than or equal to 9 under cruise conditions; and optionally wherein the ratio of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is greater than 4 and less than or equal to 4.9 under cruise conditions. [4] 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 to increase the fuel temperature upon entry into the combustion chamber (16) to at least 120°C, and preferably to at least 140°C. and optionally wherein the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) is arranged to transfer heat from the oil to the fuel to increase the fuel temperature upon entry into the combustion chamber (16) to between 120°C and 180°C, and preferably to between 140°C and 180°C. [5] Gas turbine engine (10) according to one of the preceding claims, wherein: i) the gas turbine engine further comprises at least one temperature sensor located downstream of the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706); and / or ii) the gas turbine engine is configured to receive information from a temperature sensor located in the fuel tank (50). [6] A gas turbine engine (10) according to claim 5, part i), wherein the at least one temperature sensor located downstream of the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) is located in the fuel return line (1011; 1211; 1411b; 1711b). [7] A gas turbine engine (10) according to claim 5 or claim 6, wherein the modulator valve (1010; 1210; 1410; 1710) is arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) based at least in part on temperature data from at least one temperature sensor located downstream of the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) and / or the temperature sensor located in the fuel tank (50). [8] A gas turbine engine (10) according to any preceding claim, wherein the modulator valve (1010; 1210; 1410; 1710) is arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) based at least in part on the amount of fuel remaining in the fuel tank 50. [9] A method (2300) for operating a gas turbine engine (10), the gas turbine engine (10) 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 flowed through the heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to the fuel tank (50); and a modulator valve (1010; 1210; 1410; 1710) arranged to modulate the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b); wherein the method (2300) comprises modulating (2301) the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulator valve (1010; 1210; 1410; 1710) such that a ratio of the fuel mass returned to the fuel tank (50) to the fuel mass delivered to the combustion chamber (16) is between 0 and 9 under cruise conditions. [10] The method (2300) of claim 9, wherein the method (2300) comprises modulating (2301) the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulator valve (1010; 1210; 1410; 1710) such that a ratio of the fuel mass returned to the fuel tank (50) to the fuel mass delivered to the combustion chamber (16) is between 2.3 and 9 under cruise conditions. [11] The method (2300) of claim 9 or claim 10, wherein the method (2300) comprises modulating (2301) the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulator valve (1010; 1210; 1410; 1710) such that a ratio of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is greater than 4 and less than or equal to 9 under cruise conditions, and optionally wherein the ratio of the fuel mass returned to the tank (50) to the fuel mass delivered to the combustion chamber (16) is greater than 4 and less than or equal to 4.9 under cruise conditions. [12] The method (2300) of any one of claims 9 to 11, wherein the method (2300) comprises transferring heat from the oil to the fuel using the fuel-oil heat exchanger (1004; 1006; 1204; 1206; 1404; 1406; 1704; 1706) to increase the fuel temperature upon entering the combustion chamber (16) to at least 120°C, and preferably to at least 140°C; and optionally wherein the method (2300) comprises transferring heat from the oil to the fuel using the fuel-oil heat exchanger to increase the fuel temperature upon entering the combustion chamber (16) to between 120°C and 180°C, and preferably to between 140°C and 180°C. [13] Method (2300) according to one of claims 9 to 12, wherein: i) the gas turbine engine (10) further comprises at least one temperature sensor located downstream of the heat exchanger; and / or ii) the gas turbine engine is configured to receive information from a temperature sensor located in the fuel tank (50). [14] The method (2300) of claim 13, wherein the at least one temperature sensor located downstream of the heat exchanger (which the gas turbine engine (10) includes) is located in the fuel return line (1011; 1211; 1411b; 1711b); and / or wherein the method (2300) comprises modulating the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulator valve (1010; 1210; 1410; 1710) 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. [15] The method (2300) of any one of claims 9 to 14, wherein the method (2300) comprises modulating the fuel flow along the fuel return line (1011; 1211; 1411b; 1711b) using the modulator valve (1010; 1210; 1410; 1710) based at least in part on an amount of fuel remaining in the fuel tank (50).

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