Recirculation of fuel

Abstract

A gas turbine engine for an aircraft, the gas turbine engine comprising: a combustor; a fuel-oil heat exchanger arranged to receive the fuel and to transfer heat from the oil to the fuel so as to raise the fuel temperature at the combustor inlet to at least 120 DEG C; a fuel recirculation line arranged to recirculate at least some of the fuel on the fuel flow path from a first point on the fuel flow path to a second point on the fuel flow path, the second point being upstream of the first point; and a modulation valve arranged to modulate a flow of fuel along the fuel recirculation line. A method of operating a gas turbine engine is also disclosed.

Description

[0001] Cross - reference to related applications

[0002] This specification claims the benefit of the priority of UK Patent Application No. 2319138.0, filed on December 14, 2023, the entire content of which is incorporated herein by reference. Background art Technical field

[0003] The present disclosure relates to a gas turbine engine configured to operate using a fuel different from conventional kerosene - based jet fuel, and a method of operating a gas turbine engine using a fuel different from conventional kerosene - based jet fuel.

[0004] Description of the related art

[0005] In the aviation industry, there is a trend towards using fuels different from the conventional kerosene - based jet fuels currently commonly used. It is desirable to manage the heat and fuel within the engine according to the type of fuel used. This can advantageously utilize fuel properties not possessed by kerosene - based jet fuels. Summary of the invention

[0006] According to a first aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0007] A burner;

[0008] A fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from the oil to the fuel;

[0009] A fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0010] A modulating valve 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 burner is less than 0.56.

[0011] The modulating 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 burner is less than 0.53.

[0012] The modulating 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 burner is between 0.48 and 0.56.

[0013] The modulation 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 burner (in Kelvin) is between 0.48 and 0.53.

[0014] According to a second aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0015] a burner;

[0016] a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0017] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0018] a modulation valve arranged to modulate the fuel flow along the fuel return line;

[0019] wherein the method comprises modulating the fuel flow along the fuel return line using the modulation 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 burner (in Kelvin) is less than 0.56.

[0020] The method may comprise modulating the fuel flow along the fuel return line using the modulation 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 burner (in Kelvin) is less than 0.53.

[0021] The method may comprise modulating the fuel flow along the fuel return line using the modulation 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 burner (in Kelvin) is between 0.48 and 0.56.

[0022] The method may comprise modulating the fuel flow along the fuel return line using the modulation 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 burner (in Kelvin) is between 0.48 and 0.53.

[0023] According to a third aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0024] a burner;

[0025] a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0026] A fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0027] A modulating valve 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 burner (in Kelvin) is between 0.48 and 1.00.

[0028] The modulating 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 burner (in Kelvin) is:

[0029] a) between 0.48 and 0.90;

[0030] b) between 0.48 and 0.82;

[0031] c) between 0.56 and 1.00;

[0032] d) between 0.56 and 0.82; or

[0033] e) between 0.56 and 0.75.

[0034] According to a fourth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0035] A burner;

[0036] A fuel - oil heat exchanger arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0037] A fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0038] A modulating valve arranged to modulate the fuel flow along the fuel return line;

[0039] wherein the method comprises using the modulating valve 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 burner (in Kelvin) is between 0.48 and 1.00.

[0040] The method may comprise using the modulating valve 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 burner (in Kelvin) is:

[0041] a) between 0.48 and 0.90;

[0042] b) between 0.48 and 0.82;

[0043] c) between 0.56 and 1.00;

[0044] d) between 0.56 and 0.82; or

[0045] e) between 0.56 and 0.75.

[0046] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0047] a burner;

[0048] a fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from oil to the fuel;

[0049] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0050] a modulating valve arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during start - up is less than 0.56, where ΔT is the ratio of the temperature of the fuel in the fuel tank (in Kelvin) to the temperature of the fuel delivered to the burner (in Kelvin).

[0051] The modulating valve may be arranged to modulate the fuel flow rate along the fuel return line such that the ratio of ΔT during cruise to ΔT during start - up is less than 0.53.

[0052] The modulating valve may be arranged to modulate the fuel flow rate along the fuel return line such that the ratio of ΔT during cruise to ΔT during start - up is between 0.48 and 0.56.

[0053] The modulating valve may be arranged to modulate the fuel flow rate along the fuel return line such that the ratio of ΔT during cruise to ΔT during start - up is between 0.48 and 0.53.

[0054] According to a sixth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0055] a burner;

[0056] a fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from oil to the fuel;

[0057] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0058] A modulating valve that modulates the fuel flow along a fuel return line, wherein the method includes using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is less than 0.56, where ΔT is the ratio of the temperature of the fuel in the fuel tank (in Kelvin) to the temperature of the fuel delivered to the burner (in Kelvin).

[0059] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is less than 0.53.

[0060] The method may include using the modulating valve to modulate the fuel flow rate along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is between 0.48 and 0.56.

[0061] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is between 0.48 and 0.53.

[0062] According to a seventh aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0063] A burner;

[0064] A fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from the oil to the fuel;

[0065] A fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0066] A modulating valve arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is between 0.48 and 1.88, where ΔT is the ratio of the temperature of the fuel in the fuel tank (in Kelvin) to the temperature of the fuel delivered to the burner (in Kelvin).

[0067] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT during startup is:

[0068] a) between 0.48 and 1.71;

[0069] b) between 0.48 and 1.70 (e.g., ΔT during startup is ΔT when the aircraft starts on the ground);

[0070] c) between 0.48 and 1.55 (e.g., ΔT during startup is ΔT when the aircraft starts on the ground);

[0071] d) between 0.48 and 1.88 (e.g., ΔT at startup is the ΔT when the aircraft starts up in the air); or

[0072] e) between 0.48 and 1.71 (where ΔT at startup is the ΔT when the aircraft starts up in the air).

[0073] According to an eighth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0074] a burner;

[0075] a fuel - oil heat exchanger arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0076] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0077] a modulating valve arranged to modulate the fuel flow along the fuel return line, wherein the method comprises using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT at startup is between 0.48 and 1.88, where ΔT is the ratio of the temperature of the fuel in the fuel tank (in Kelvin) to the temperature of the fuel delivered to the burner (in Kelvin).

[0078] The method may comprise using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of ΔT during cruise to ΔT at startup is:

[0079] a) between 0.48 and 1.71;

[0080] b) between 0.48 and 1.70 (e.g., ΔT at startup is the ΔT when the aircraft starts up on the ground);

[0081] c) between 0.48 and 1.55 (e.g., ΔT at startup is the ΔT when the aircraft starts up on the ground);

[0082] d) between 0.48 and 1.88 (e.g., ΔT at startup is the ΔT when the aircraft starts up in the air); or

[0083] e) between 0.48 and 1.71 (where ΔT at startup is the ΔT when the aircraft starts up in the air).

[0084] According to a ninth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0085] a burner;

[0086] A fuel - oil heat exchanger, which is arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0087] A fuel return line, which is arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0088] A modulating valve, which is arranged to modulate the fuel flow along the fuel return line, wherein when the fuel that has passed through the heat exchanger is at a temperature of at least 120 °C, the modulating valve is arranged to initiate returning the fuel to the fuel tank.

[0089] The modulating valve can be arranged to initiate returning the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature of at least 140 °C.

[0090] The modulating valve can be arranged to initiate returning the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature between 120 °C and 180 °C, and preferably between 140 °C and 180 °C.

[0091] The modulating valve can be arranged to initiate returning the fuel to the fuel tank in the following cases:

[0092] i) Provide an indication of the operating conditions; and

[0093] ii) The temperature of the fuel that has passed through the heat exchanger is at least 120 °C, preferably at least 140 °C.

[0094] The operating conditions can be one or more of the following: the proportion of sustainable aviation fuel (SAF) in the fuel, the thermal stability of the fuel, the coking level of the fuel, the oxygen content of the fuel, and the sulfur content of the fuel.

[0095] The gas turbine engine may further include a sensor configured to detect one or more operating conditions.

[0096] According to a tenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0097] A burner;

[0098] A fuel - oil heat exchanger, which is arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0099] A fuel return line, which is arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0100] A modulating valve, which is arranged to modulate the fuel flow along the fuel return line;

[0101] Wherein, the method includes using a modulating valve to initiate returning the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature of at least 120°C.

[0102] The method may include using a modulating valve to initiate returning the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature of at least 140°C.

[0103] The method may include using a modulating valve to initiate returning the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature between 120°C and 180°C, and preferably between 140°C and 180°C.

[0104] The method may include using a modulating valve to initiate returning the fuel to the fuel tank when:

[0105] i) providing an indication of operating conditions; and

[0106] ii) the temperature of the fuel that has passed through the heat exchanger is at least 120°C, preferably at least 140°C.

[0107] The operating conditions may be one or more of the following: the proportion of sustainable aviation fuel (SAF) in the fuel, the thermal stability of the fuel, the coking level of the fuel, the oxygen content of the fuel, and the sulfur content of the fuel.

[0108] The gas turbine engine may further include a sensor configured to detect one or more operating conditions. The method may include using the sensor to detect one or more operating conditions.

[0109] According to the eleventh aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0110] a combustor;

[0111] a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from the oil to the fuel;

[0112] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0113] a modulating valve arranged to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the combustor is between 0 and 9 under cruise conditions.

[0114] "The fuel mass returned to the fuel tank" and "the fuel mass delivered to the burner" can be expressed as fuel flow rates of the fuel mass in the fuel flow per unit time (e.g., in kg / s). The "fuel mass delivered to the burner" in the eleventh and twelfth aspects refers to the "fuel mass burned", e.g., the amount of fuel burned by the burner per unit time of fuel flow rate.

[0115] The modulation valve can 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 burner under cruise conditions is between 2.3 and 9.

[0116] The modulation valve can 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 burner under cruise conditions is greater than 4 and less than or equal to 9.

[0117] The modulation valve can 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 burner under cruise conditions is greater than 4 and less than or equal to 4.9.

[0118] According to the twelfth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[0119] A burner;

[0120] A fuel-oil heat exchanger arranged to receive fuel from a fuel tank on an aircraft and transfer heat from oil to the fuel;

[0121] A fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0122] A modulation valve arranged to modulate the fuel flow along the fuel return line;

[0123] wherein the method comprises modulating the fuel flow along the fuel return line using the modulation valve such that the ratio of the fuel mass returned to the fuel tank to the fuel mass delivered to the burner under cruise conditions is between 0 and 9.

[0124] The method may comprise modulating the fuel flow along the fuel return line using the modulation valve such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the burner under cruise conditions is between 2.3 and 9.

[0125] The method may comprise modulating the fuel flow along the fuel return line using the modulation valve such that the ratio of the fuel mass returned to the tank to the fuel mass delivered to the burner under cruise conditions is greater than 4 and less than or equal to 9.

[0126] The method may include modulating a fuel flow along a fuel return line using a modulating valve such that, under cruise conditions, the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 4.9.

[0127] According to a thirteenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0128] a burner;

[0129] a fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from oil to the fuel;

[0130] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0131] a modulating valve arranged to modulate the fuel flow along the fuel return line, wherein the modulating valve is arranged to prevent fuel at a temperature of 180 °C or higher from returning to the fuel tank.

[0132] The modulating valve may be arranged to prevent fuel from returning to the fuel tank when the fuel temperature in the fuel tank is at a predetermined upper threshold temperature.

[0133] 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.

[0134] The modulating valve may be arranged to modulate the fuel flow along the fuel return line at least partially based on the temperature of the fuel that has passed through the heat exchanger, the temperature of the fuel in the fuel tank, and the amount of fuel remaining in the tank.

[0135] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that once the fuel in the fuel tank is mixed with the fuel returning to the tank along the fuel return line, the equilibrium temperature of the fuel does not exceed a predetermined upper threshold temperature.

[0136] According to a fourteenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0137] a burner;

[0138] a fuel - oil heat exchanger arranged to receive fuel from a fuel tank on the aircraft and transfer heat from oil to the fuel;

[0139] a fuel return line arranged to return at least some of the fuel that has passed through the heat exchanger to the fuel tank; and

[0140] A modulating valve, which is arranged to modulate the fuel flow along a fuel return line;

[0141] Wherein, the method includes using the modulating valve to prevent fuel at a temperature of 180 °C or higher from returning to the fuel tank.

[0142] The method may include using the modulating valve to prevent fuel from returning to the fuel tank when the fuel temperature in the fuel tank is at a predetermined upper threshold temperature.

[0143] 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.

[0144] The method may include using the modulating valve to modulate the fuel flow along the fuel return line at least partially based on the temperature of the fuel that has passed through the heat exchanger, the temperature of the fuel in the fuel tank, and the amount of fuel remaining in the tank.

[0145] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that once the fuel in the fuel tank is mixed with the fuel returning to the tank along the fuel return line, the equilibrium temperature of the fuel does not exceed the predetermined upper threshold temperature.

[0146] According to the fifteenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0147] A burner;

[0148] A fuel-oil heat exchanger, which is arranged to receive fuel and transfer heat from oil to the fuel so as to raise the fuel temperature at the burner inlet to at least 120 °C;

[0149] A fuel recirculation line, which is arranged to recirculate at least some of the fuel on the fuel flow path from a first point on the fuel flow path to a second point on the fuel flow path, the second point being upstream of the first point; and

[0150] A modulating valve, which is arranged to modulate the fuel flow along the fuel recirculation line.

[0151] 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.

[0152] The gas turbine engine may further include a fuel pump located on 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.

[0153] The fuel pump may be located downstream of the fuel-oil heat exchanger in the fuel flow path.

[0154] The fuel recirculation line may be arranged to supply fuel to one or more additional aircraft and / or engine mechanisms.

[0155] One or more additional aircraft and / or engine mechanisms may be or include one or more of an nacelle anti-icing system, an actuator, a bleed valve, a thermal management modulation valve (e.g., for engine and / or generator system and turbine housing cooling systems).

[0156] The fuel flow path may be the main fuel flow path between the fuel tank and the burner on the aircraft. The fuel-oil heat exchanger may be the primary fuel-oil heat exchanger located on the main fuel flow path.

[0157] The second point may be located downstream of the low-pressure fuel pump through which fuel is pumped from the fuel tank to the gas turbine engine.

[0158] The first point may be located downstream of the primary fuel-oil heat exchanger. The first point may be located downstream of the fuel pump (i.e., the high-pressure fuel pump) configured to supply fuel to the burner.

[0159] The engine may further include a secondary fuel-oil heat exchanger (e.g., in series with the main heat exchanger) in the main flow path. The second point may be downstream of the secondary heat exchanger.

[0160] Alternatively, the engine may further include a secondary fuel-oil heat exchanger having a flow line leading from the main flow path to the secondary fuel-oil heat exchanger and a flow line rejoining the main flow path from the secondary heat exchanger (e.g., such that it is parallel to the main flow path). The second point may be upstream of the line inlet leading to the secondary fuel-oil heat exchanger, and the first point may be downstream of the line outlet rejoining the main fuel flow path from the secondary fuel-oil heat exchanger. The second point may be upstream of the primary fuel-oil heat exchanger.

[0161] The fuel-oil heat exchanger may be arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to at least 140 °C (i.e., the heat exchanger may be arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature to these temperatures before the fuel enters the burner).

[0162] The heat exchanger may be arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to between 120 °C and 180 °C, and preferably to between 140 °C and 180 °C.

[0163] The heat exchanger may be arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet 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 raise the fuel temperature to a value defined within the range of any two of these values.

[0164] The modulation valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the mass of the recirculated fuel to the mass of the fuel delivered to the burner is between 0 and 9 under cruise conditions.

[0165] The modulation valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the mass of the fuel returned to the tank to the mass of the fuel delivered to the burner is between 2.3 and 9 under cruise conditions.

[0166] The modulation valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the mass of the fuel returned to the tank to the mass of the fuel delivered to the burner is greater than 4 and less than or equal to 9 under cruise conditions.

[0167] The modulation valve may be arranged to modulate the fuel flow along the fuel recirculation line such that the ratio of the mass of the fuel returned to the tank to the mass of the fuel delivered to the burner is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0168] The gas turbine engine may further include at least one temperature sensor in the fuel flow path and / or in the fuel recirculation line.

[0169] The modulation valve may be arranged to modulate the fuel flow along the fuel recirculation line at least in part based on temperature data from the at least one temperature sensor.

[0170] According to a sixteenth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0171] a burner;

[0172] a fuel-oil heat exchanger arranged to receive fuel from a fuel tank and to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to at least 120 °C;

[0173] a fuel recirculation line arranged to recirculate at least some of the fuel in the fuel flow path from a first point on the fuel flow path to a second point on the fuel flow path, the second point being upstream of the first point; and

[0174] A modulation valve, which is arranged to modulate the fuel flow along a fuel recirculation line;

[0175] Wherein, the method includes using the modulation valve to modulate the fuel flow along the fuel recirculation line.

[0176] The method may include using the modulation valve to modulate the fuel flow along the fuel recirculation line from a first point downstream of a fuel-oil heat exchanger to a second point upstream of the fuel-oil heat exchanger.

[0177] The gas turbine engine may further include a fuel pump located on the fuel flow path. The method may include using the modulation valve to modulate the fuel flow along the fuel recirculation line from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

[0178] The fuel pump may be located downstream of a fuel-oil heat exchanger on the fuel flow path.

[0179] The method may include supplying fuel to one or more additional aircraft and / or engine mechanisms via the fuel recirculation line.

[0180] One or more additional aircraft and / or engine mechanisms may be or include one or more of a nacelle anti-icing system, an actuator, a bleed valve, a thermal management modulation valve (e.g., for an engine and / or generator system and a turbine casing cooling system).

[0181] The fuel flow path may be the main fuel flow path between a fuel tank on the aircraft and a burner. The fuel-oil heat exchanger may be a primary fuel-oil heat exchanger located on the main fuel flow path.

[0182] The second point may be located downstream of a low-pressure fuel pump through which fuel is pumped from the fuel tank to the gas turbine engine.

[0183] 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 (i.e., a high-pressure fuel pump) configured to supply fuel to the burner.

[0184] The engine may further include a secondary fuel-oil heat exchanger (e.g., in series with the main heat exchanger) located in the main flow path. The second point may be downstream of the secondary heat exchanger.

[0185] Alternatively, the engine may further include a secondary fuel - oil heat exchanger having a flow line leading from the main flow path to the secondary fuel - oil heat exchanger and a flow line rejoining the main flow path from the secondary heat exchanger (e.g., such that it is parallel to the main flow path). The second point may be upstream of the line inlet leading to the secondary fuel - oil heat exchanger, and the first point may be downstream of the line outlet rejoining the main fuel flow path from the secondary fuel - oil heat exchanger. The second point may be upstream of the primary fuel - oil heat exchanger.

[0186] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to at least 140 °C.

[0187] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to a temperature between 120 °C and 180 °C.

[0188] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to a temperature between 140 °C and 180 °C.

[0189] The heat exchanger may be arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet 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 raise the fuel temperature to a temperature defined within the range of any two of these values.

[0190] The method may include modulating the fuel flow along the recirculation line using a modulating valve such that the ratio of the mass of fuel recirculated under cruise conditions to the mass of fuel delivered to the burner is between 0 and 9.

[0191] The method may include modulating the fuel flow along the recirculation line using a modulating valve such that the ratio of the mass of fuel returned to the tank under cruise conditions to the mass of fuel delivered to the burner is between 2.3 and 9.

[0192] The method may include modulating the fuel flow along the recirculation line using a modulating valve such that the ratio of the mass of fuel returned to the tank under cruise conditions to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 9.

[0193] The method may include modulating the fuel flow along the recirculation line using a modulating valve such that the ratio of the mass of fuel returned to the tank under cruise conditions to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 4.9.

[0194] The gas turbine engine may also include at least one temperature sensor in the fuel flow path and / or the fuel recirculation line.

[0195] The method may include modulating the fuel flow along the fuel recirculation line using a modulating valve based at least in part on temperature data from at least one temperature sensor.

[0196] The fuel recirculation features of the fifteenth and sixteenth aspects may be combined with any other aspect defined above or elsewhere herein. In other words, a modulating valve of any aspect defined above or elsewhere herein may be used to recirculate fuel back to the fuel tank and recirculate as defined in the fifteenth and sixteenth aspects.

[0197] In the fifth, seventh, ninth, eleventh, and thirteenth aspects :

[0198] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is less than 0.56.

[0199] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is less than 0.53.

[0200] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 0.56.

[0201] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 0.53.

[0202] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 1.00.

[0203] The modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is:

[0204] a) between 0.48 and 0.90;

[0205] b) between 0.48 and 0.82;

[0206] c) between 0.56 and 1.00;

[0207] d) between 0.56 and 0.82; or

[0208] e) between 0.56 and 0.75.

[0209] In the sixth, eighth, tenth, twelfth, and fourteenth aspects :

[0210] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is less than 0.56.

[0211] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is less than 0.53.

[0212] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 0.56.

[0213] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 0.53.

[0214] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is between 0.48 and 1.00.

[0215] The method may include modulating the fuel flow along the fuel return line using a modulating 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 burner (in Kelvin) (e.g., under cruise conditions) is:

[0216] a) between 0.48 and 0.90;

[0217] b) between 0.48 and 0.82;

[0218] c) between 0.56 and 1.00;

[0219] d) between 0.56 and 0.82; or

[0220] e) between 0.56 and 0.75.

[0221] In the first, third, fifth, seventh, eleventh, and thirteenth aspects :

[0222] The modulating valve can be arranged to initiate the return of the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature of at least 120 °C and preferably at least 140 °C.

[0223] The modulating valve can be arranged to initiate the return of the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature between 120 °C and 180 °C and preferably between 140 °C and 180 °C.

[0224] The modulating valve can be arranged to initiate the return of the fuel to the fuel tank when:

[0225] i) an indication of the operating conditions is provided; and

[0226] ii) the temperature of the fuel that has passed through the heat exchanger is at least 120 °C, preferably at least 140 °C.

[0227] The operating conditions can be one or more of the following: the proportion of sustainable aviation fuel (SAF) in the fuel, the thermal stability of the fuel, the coking level of the fuel, the oxygen content of the fuel, and the sulfur content of the fuel.

[0228] The gas turbine engine can also include a sensor configured to detect one or more operating conditions.

[0229] In the second, fourth, sixth, eighth, twelfth, and fourteenth aspects :

[0230] The method can include using the modulating valve to initiate the return of the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature of at least 120 °C and preferably at least 140 °C.

[0231] The method can include using the modulating valve to initiate the return of the fuel to the fuel tank when the fuel that has passed through the heat exchanger is at a temperature between 120 °C and 180 °C, preferably between 140 °C and 180 °C.

[0232] The method can include using the modulating valve to initiate the return of the fuel to the fuel tank when:

[0233] i) an indication of the operating conditions is provided; and

[0234] ii) The temperature of the fuel that has passed through the heat exchanger is at least 120 °C, preferably at least 140 °C.

[0235] The operating conditions can be one or more of the following: the proportion of sustainable aviation fuel (SAF) in the fuel, the thermal stability of the fuel, the coking level of the fuel, the oxygen content of the fuel, and the sulfur content of the fuel.

[0236] The gas turbine engine may further include a sensor configured to detect one or more operating conditions. The method may include using the sensor to detect one or more operating conditions.

[0237] In the first, third, fifth, seventh, ninth, and thirteenth aspects

[0238] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is between 0 and 9 under cruise conditions.

[0239] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is between 2.3 and 9 under cruise conditions.

[0240] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 9 under cruise conditions.

[0241] The modulating valve may be arranged to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0242] In the second, fourth, sixth, eighth, tenth, and fourteenth aspects :

[0243] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the fuel tank to the mass of fuel delivered to the burner is between 0 and 9 under cruise conditions.

[0244] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is between 2.3 and 9 under cruise conditions.

[0245] The method may include using the modulating valve to modulate the fuel flow along the fuel return line such that the ratio of the mass of fuel returned to the tank to the mass of fuel delivered to the burner is greater than 4 and less than or equal to 9 under cruise conditions.

[0246] The method may include modulating a fuel flow along a fuel return line using a modulating valve such that a ratio of a fuel mass returned to a tank to a fuel mass delivered to a burner is greater than 4 and less than or equal to 4.9 under cruise conditions.

[0247] In the first, third, fifth, seventh, ninth, and eleventh aspects :

[0248] The modulating valve may be arranged to prevent fuel at a temperature of 180 °C or higher from returning to the fuel tank.

[0249] The modulating valve may be arranged to prevent fuel from returning to the fuel tank when the fuel temperature in the fuel tank is at a predetermined upper threshold temperature.

[0250] 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.

[0251] The modulating valve may be arranged to modulate the fuel flow along the fuel return line at least in part based on a temperature of fuel that has passed through a heat exchanger, a temperature of fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0252] The modulating 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 does not exceed a predetermined upper threshold temperature once the fuel in the fuel tank is mixed with the fuel returning to the tank along the fuel return line.

[0253] In the second, fourth, sixth, eighth, tenth, and twelfth aspects :

[0254] The method may include using a modulating valve to prevent fuel at a temperature of 180 °C or higher from returning to the fuel tank.

[0255] The method may include using a modulating valve to prevent fuel from returning to the fuel tank when the fuel temperature in the fuel tank is at a predetermined upper threshold temperature.

[0256] 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.

[0257] The method may include using a modulating valve to modulate the fuel flow along the fuel return line at least in part based on a temperature of fuel that has passed through a heat exchanger, a temperature of fuel in the fuel tank, and an amount of fuel remaining in the tank.

[0258] The method may include modulating the fuel flow along the fuel return line using a modulating valve such that once the fuel in the fuel tank is mixed with the fuel returning to the tank along the fuel return line, the equilibrium temperature of the fuel does not exceed a predetermined upper threshold temperature.

[0259] In the first, second, third, and fourth aspects :

[0260] The temperature of the fuel delivered to the burner (i.e., at the burner inlet) may be the temperature of the fuel delivered to the burner under cruise conditions. The fuel temperature delivered to the burner under cruise conditions may be defined as the average value over at least 5 minutes, 10 minutes, or 30 minutes under steady-state cruise conditions. These average temperatures do not include transient spikes in temperature, which may be defined as fluctuations in the fuel temperature during operation, typically an increase in temperature. Each fluctuation may last no more than 5 minutes.

[0261] The previous paragraph above may also be related to other references of the fuel temperature at the burner inlet.

[0262] In the first, third, fifth, seventh, ninth, eleventh, thirteenth and fifteenth aspects :

[0263] The heat exchanger may be arranged to transfer heat from the oil to the fuel in order to raise the fuel temperature at the burner inlet 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 to raise the temperature of the fuel such that the fuel has those temperatures when entering the burner).

[0264] The heat exchanger may be arranged to transfer heat from the oil to the fuel in order to raise the fuel temperature at the burner inlet to between 120 °C and 180 °C and preferably to between 140 °C and 180 °C.

[0265] The heat exchanger may be arranged to transfer heat from the oil to the fuel in order to raise the fuel temperature at the burner inlet 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 raise the fuel temperature defined within the range of any two of these values.

[0266] In the second, fourth, sixth, eighth, tenth, twelfth, and fourteenth aspects and sixteenth aspects :

[0267] The method may include using a fuel-oil heat exchanger to transfer heat from the oil to the fuel in order to raise the fuel temperature at the burner inlet to at least 120 °C and preferably to at least 140 °C.

[0268] The method may include using a fuel - oil heat exchanger to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to a temperature between 120°C and 180°C, and preferably to a temperature between 140°C and 180°C.

[0269] The method may include transferring heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet 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 defined within a range between any two of these values.

[0270] In the first, third, fifth, seventh aspects :

[0271] The gas turbine engine may include a first temperature sensor downstream of the heat exchanger; and may be configured to receive information from a second temperature sensor located in the fuel tank.

[0272] The first temperature sensor may be located in the fuel return line.

[0273] The modulating 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.

[0274] In the ninth aspect :

[0275] The gas turbine engine may further include at least one temperature sensor downstream of the heat exchanger.

[0276] The modulating valve may be arranged to initiate returning fuel to the fuel tank at least in part based on temperature data from at least one temperature sensor downstream of the heat exchanger.

[0277] In the eleventh and thirteenth aspects :

[0278] The gas turbine engine may further include at least one temperature sensor 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.

[0279] At least one temperature sensor downstream of the heat exchanger may be located in the fuel return line.

[0280] The modulating valve may be arranged to modulate the fuel flow along the fuel return line at least in part based on temperature data from at least one temperature sensor downstream of the heat exchanger and / or from a temperature sensor located in the fuel tank.

[0281] In the second, fourth, sixth, eighth aspects :

[0282] A gas turbine engine may include a first temperature sensor downstream of a heat exchanger; and may be configured to receive information from a second temperature sensor located in a fuel tank.

[0283] The first temperature sensor may be located in a fuel return line.

[0284] The method may include modulating a fuel flow along the fuel return line based on temperature data from the first temperature sensor and the second temperature sensor.

[0285] In the tenth aspect :

[0286] The gas turbine engine may further include at least one temperature sensor downstream of the heat exchanger.

[0287] The method may include initiating a return of fuel to the fuel tank using a modulating valve based at least in part on temperature data from at least one temperature sensor downstream of the heat exchanger.

[0288] In the twelfth and fourteenth aspects :

[0289] The gas turbine engine may further include at least one temperature sensor 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.

[0290] At least one temperature sensor downstream of the heat exchanger may be located in the fuel return line.

[0291] The method may include modulating a fuel flow along the fuel return line using a modulating valve based at least in part on temperature data from at least one temperature sensor downstream of the heat exchanger and / or a temperature sensor located in the fuel tank.

[0292] In the first, second, third, fourth, fifth, sixth, seventh, eighth aspects, ninth, tenth, eleventh, twelfth, thirteenth, fourteenth aspects :

[0293] The temperature of the fuel in the fuel tank may be between -54°C and 100°C.

[0294] The temperature of the fuel in the fuel tank may be between -54°C and 65°C.

[0295] The temperature of the fuel in the fuel tank may be between -54°C and 55°C.

[0296] The temperature of the fuel in the fuel tank may be between -54°C and 5°C.

[0297] 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 be limited to any range between any two of these values.

[0298] In the first, third, fifth, seventh, ninth, eleventh, thirteenth aspects :

[0299] The modulation valve can be arranged to modulate the fuel flow along the fuel return line at least partially based on the amount of fuel remaining in the fuel tank. The amount of fuel remaining in the fuel tank can be the mass of the fuel remaining in the fuel tank.

[0300] The modulation valve can be arranged to modulate the fuel flow along the fuel return line at least partially based on the temperature of the fuel that has passed through the heat exchanger, the temperature of the fuel in the fuel tank, and the amount of fuel remaining in the fuel tank.

[0301] The modulation valve can be arranged to modulate the fuel flow along the fuel return line such that after the fuel returns to the fuel tank along the fuel return line, the equilibrium temperature of the fuel in the fuel tank does not exceed a predetermined upper threshold temperature of the fuel in the fuel tank.

[0302] In the second, fourth, sixth, eighth, tenth, twelfth, fourteenth aspects :

[0303] The method can include modulating the fuel flow along the fuel return line using the modulation valve at least partially based on the amount of fuel remaining in the fuel tank. The amount of fuel remaining in the fuel tank can be the mass of the fuel remaining in the fuel tank.

[0304] The method can include modulating the fuel flow along the fuel return line using the modulation valve at least partially based on the temperature of the fuel that has passed through the heat exchanger, the temperature of the fuel in the fuel tank, and the amount of fuel remaining in the fuel tank.

[0305] The method can include modulating the fuel flow along the fuel return line using the modulation valve such that after the fuel returns to the fuel tank along the fuel return line, the equilibrium temperature of the fuel in the fuel tank does not exceed a predetermined upper threshold temperature of the fuel in the fuel tank.

[0306] As described elsewhere herein, the present disclosure may be applied to any relevant configuration of a gas turbine engine. Such gas turbine engines may be, for example, turbofan gas turbine engines, open rotor gas turbine engines (where the propellers are not enclosed by nacelles), turboprop engines or turbojet engines. Any such engine may or may not be provided with a afterburner. Such gas turbine engines may be configured, for example, for land or marine power generation applications.

[0307] A gas turbine engine according to any aspect of the present disclosure may include an engine core that includes a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor. Such gas turbine engines may include a fan (with fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, such as in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propulsor).

[0308] In the case where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two contra-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propulsors may rotate in opposite directions such that one propulsor rotates clockwise about the axis of rotation of the engine and the other propulsor rotates counterclockwise about the axis of rotation of the engine. Alternatively, the gas turbine engine may include a propulsor stage and a stator vane stage constructed downstream of the propulsor stage. The stator vane stage may have variable pitch. Thus, the high pressure, intermediate pressure, and free power turbines may drive the high pressure and intermediate pressure compressors and the propulsors respectively via suitable interconnecting shafts. Thus, the propulsors may provide most of the propulsive thrust.

[0309] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more propulsor stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0310] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct drive turbofan engine where the fan is directly connected to a fan drive turbine via a spool, for example without a gearbox. In such a direct drive turbofan engine, it can be said that the fan rotates at the same rotational speed as the fan drive turbine. By way of example only, the fan drive turbine may be a first turbine, the spool may be a first spool, and the gas turbine engine may further include a second turbine and a second spool connecting the second turbine to the compressor. The second turbine, compressor, and second spool may be arranged to rotate at a higher rotational speed than the first spool. In such an arrangement, the second turbine may be axially located upstream of the first turbine.

[0311] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan driven via a gearbox. Thus, such a gas turbine engine may include a gearbox that receives an input from a spool and outputs drive to the fan so as to drive the fan at a lower rotational speed than the spool. The input to the gearbox may come directly from the spool or indirectly from the spool, such as via spur shafts and / or gears. The spool may rigidly connect the turbine and the compressor such that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0312] A gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and the compressor, such as one shaft, two shafts, or three shafts. By way of example only, the turbine connected to the spool may be a first turbine, the compressor connected to the spool may be a first compressor, and the spool may be a first spool. The engine core may also include a second turbine, a second compressor, and a second spool connecting the second turbine to the second compressor. The second turbine, second compressor, and second spool may be arranged to rotate at a higher rotational speed than the first spool.

[0313] In such an arrangement, the second compressor may be axially located downstream of the first compressor. The second compressor may be arranged to receive flow (e.g., directly receive, such as via a substantially annular duct) from the first compressor.

[0314] The gearbox may be arranged to be driven by a spool (such as the first spool in the above example) configured (e.g., in use) to rotate at the lowest rotational speed. For example, the gearbox may be arranged to be driven only by a spool configured (e.g., in use) to rotate at the lowest rotational speed (e.g., in the above example, only by the first spool and not the second spool). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first shaft and / or the second shaft in the above example.

[0315] The gearbox may be a reduction gearbox (since the output to the fan has a lower rotational rate than the input from the spool). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. Such a gearbox may be single stage. Alternatively, such a gearbox may be a compound gearbox, such as a compound planetary gearbox (which may have an input on the sun gear and an output on the ring gear and is thus referred to as a "compound star" gearbox), such as having two stages of reduction.

[0316] The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), e.g., greater than 2.5, e.g., in the range of 3 to 4.2, or 3.2 to 3.8, e.g., approximately 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. For example, the gear ratio can be between any two values in the previous sentence. By way of example only, the gearbox can be a "star" gearbox having a reduction ratio in the range of 3.1 or 3.2 to 3.8. By way of another example only, the gearbox can be a "star" gearbox having a reduction ratio in the range of 3.0 to 3.1. By way of yet another example only, the gearbox can be a "planetary" gearbox having a reduction ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio can be outside of these ranges.

[0317] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or blend is provided to a combustor that can be disposed downstream (e.g., axially downstream) of the fan and compressor with respect to the flow path. For example, in the case where a second compressor is provided, the combustor can be located directly downstream of the second compressor (e.g., at its outlet). By way of another example, in the case where a second turbine is provided, the flow at the combustor outlet can be provided to the inlet of the second turbine. The combustor can be disposed upstream of one or more turbines.

[0318] The compressor or each compressor (e.g., the first compressor and the second compressor as described above) can include any number of stages, e.g., a plurality of stages. Each stage can include a row of rotor blades and a row of stator vanes, and the row of stator vanes can be variable stator vanes (since the angle of incidence of the row of stator vanes can be variable). The row of rotor blades and the row of stator vanes can be axially offset from each other. For example, the gas turbine engine can be a direct drive turbofan gas turbine engine including 13 or 14 compressor stages (in addition to the fan). Such an engine can include, for example, 3 stages in the first (or "low pressure") compressor and 10 or 11 stages in the second (or "high pressure") compressor. By way of another example, the gas turbine engine can be a "geared" gas turbine engine including 11, 12, or 13 compressor stages (in addition to the fan) (where the fan is driven by a first shaft via a reduction gearbox). Such an engine can include 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. By way of yet another example, the gas turbine engine can be a "geared" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0319] The turbine or each turbine (e.g., the first and second turbines as described above) may include any number of stages, such as a plurality of stages. Optionally, each stage may include a row of rotor blades and a row of stator vanes, and vice versa. The corresponding rows of rotor blades and stator vanes may be axially offset from each other. The second (or "high pressure") turbine may include 2 stages in any arrangement (e.g., whether it is a geared engine or a direct drive engine). The gas turbine engine may be a direct drive gas turbine engine including a first (or "low pressure") turbine having 5, 6 or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine including a first (or "low pressure") turbine having 3 or 4 stages.

[0320] Each fan blade may be defined as having a radial span that extends from a root (or hub) at a radially inner gas washing position or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or approximately) any of the following: 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 the inclusive range defined by any two values in the previous sentence (i.e., these values may form an upper or lower limit), e.g., in the range of 0.28 to 0.32 or 0.29 to 0.30. These ratios are generally referred to as hub-tip ratios. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially foremost) portion of the blade. Of course, the hub-tip ratio refers to the gas washing portion of the fan blade, i.e., the portion radially outside any platform.

[0321] The radius of the fan can be measured between the centerline of the engine and the tip at the leading edge of the fan blade. The fan diameter (which may simply be twice the fan radius) can be greater than (or approximately) any of the following: 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 within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form the upper or lower limit), such as in the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. By way of non-limiting example only, the fan diameter can be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0322] The rotational speed of the fan can vary during use. Generally speaking, for fans with a larger diameter, the rotational speed is lower. By way of non-limiting examples only, the rotational speed of the fan under cruise conditions can be less than 3500 rpm, such as less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 200 cm to 210 cm, the rotational speed of the fan under cruise conditions can be in the range of 2750 rpm to 2900 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 210 cm to 230 cm, the rotational speed of the fan under cruise conditions can be in the range of 2500 rpm to 2800 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 340 cm to 360 cm, the rotational speed of the fan under cruise conditions can be in the range of 1500 rpm to 1800 rpm. By way of additional non-limiting examples only, for a direct drive engine with a fan diameter in the range of 190 cm to 200 cm, the rotational speed of the fan under cruise conditions can be in the range of 3600 rpm to 3900 rpm. By way of additional non-limiting examples only, for a direct drive engine with a fan diameter in the range of 300 cm to 340 cm, the rotational speed of the fan under cruise conditions can be in the range of 2000 rpm to 2800 rpm.

[0323] When using a gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation causes the tips of the fan blades to move at a speed U 尖端 The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. The fan tip loading can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D mean enthalpy rise), and U 尖端 is the (translational) speed of the fan tip, e.g., at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip loading under cruise conditions can be greater than (or approximately) any of the following: 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 are dimensionless). The fan tip loading can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), e.g., in the range of 0.28 to 0.31 or 0.29 to 0.3 (e.g., for a geared gas turbine engine).

[0324] A gas turbine engine according to the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio at cruise conditions may be greater than (or approximately) any of the following: 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 12 to 16, or 13 to 15, or 13 to 14. By way of non-limiting example only, the bypass ratio at 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 another non-limiting example only, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be at least substantially annular. The bypass duct may be located radially outward of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0325] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the exit of the highest pressure compressor (before entering the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 50 to 70. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0326] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, for a given thrust condition, the specific thrust can depend on the particular composition of the fuel provided to the combustor. Under cruise conditions, the specific thrust of the engines described and / or claimed herein can be less than (or approximately) any of the following: 110 N / kg -1 s, 105 N / kg -1 s, 100 N / kg -1 s, 95 N / kg -1 s, 90 N / kg -1 s, 85 N / kg -1 s or 80 N / kg -1 s. The specific thrust can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, between 80 N / kg -1 s and 100 N / kg -1 s, or between 85 N / kg -1 s and 95 N / kg -1 s. Compared to conventional gas turbine engines, such engines can be particularly efficient. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be between 90 N / kg -1 s and 95 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be between 80 N / kg -1 s and 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be between 70 N / kg -1 s and 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be between 90 N / kg -1 s and 120 N / kg -1 s.

[0327] A gas turbine engine as described herein and / or claimed may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described herein and / or claimed may produce a maximum thrust of at least (or approximately) any one 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 defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit). By way of non-limiting example only, a gas turbine as described herein and / or claimed may be capable of producing a maximum thrust within 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, a geared gas turbine engine having a fan diameter within the range of 200 cm to 210 cm may have a maximum thrust within the range of 140 kN to 160 kN. By way of non-limiting example only, a geared gas turbine engine having a fan diameter within the range of 210 cm to 230 cm may have a maximum thrust within the range of 150 kN to 200 kN. By way of non-limiting example only, a geared gas turbine engine having a fan diameter within the range of 340 cm to 360 cm may have a maximum thrust within the range of 370 kN to 500 kN. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter within the range of 300 cm to 340 cm may have a maximum thrust within the range of 370 kN to 500 kN. The thrust mentioned above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15 °C (ambient pressure 101.3 kPa, temperature 30 °C), with the engine stationary.

[0328] In use, the temperature of the flow at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be referred to as TET, can be measured at the outlet of the combustor, for example, just upstream of the first turbine blade which may itself be referred to as the nozzle guide vane. In some examples, for a given thrust condition, the TET can depend on the particular composition of the fuel supplied to the combustor. Under cruise conditions, the TET can be at least (or approximately) any of the following: 1400K, 1450K, 1500K, 1520K, 1530K, 1540K, 1550K, 1600K or 1650K. The TET under cruise conditions can be within the range bounded by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example 1530K to 1600K. Thus, by way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can have a TET under cruise conditions in the range of 1540K to 1600K. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can have a TET under cruise conditions in the range of 1590K to 1650K. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can have a TET under cruise conditions in the range of 1600K to 1660K. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET under cruise conditions in the range of 1590K to 1650K. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET under cruise conditions in the range of 1570K to 1630K.

[0329] The maximum TET during engine operation can be, for example, at least (or approximately) any one of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, 2000K, 2050K, or 2100K. The maximum TET can be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 1800K to 1950K or 1900K to 2000K. Thus, by way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1935K to 1995K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1890K to 1950K. The maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0330] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein can be made of any suitable material or combination of materials. For example, at least a portion of the fan blades and / or the airfoil can be made at least in part of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. By way of further example, at least a portion of the fan blades and / or the airfoil can be made at least in part of 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 blades can include at least two regions made of different materials. For example, the fan blades can have a protective leading edge that can be made of a material that better resists impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such leading edges can be made, for example, of titanium or a titanium-based alloy. Thus, by way of example only, the fan blade can have a carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0331] The fan as described herein and / or claimed may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixture that may engage a corresponding slot in a hub (or disc portion). By way of example only, such a fixture may be in the form of a dovetail that may be inserted into and / or engage a corresponding slot in the hub / disc portion so as to secure the fan blade to the hub / disc portion. By way of a further example, the fan blades may be integrally formed with the central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc portion by welding (such as linear friction welding).

[0332] The gas turbine engine described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to vary in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0333] The fan of a gas turbine as described herein and / or claimed may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades. In the case where the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. In the case where the fan blades have a metal body (such as an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0334] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, landing, approach, and landing (or one or more portions thereof) have their conventional meanings and will be readily understood by the person skilled in the art. Thus, for a given gas turbine engine for an aircraft, the person skilled in the art will immediately recognize that each term is used to refer to the overall or one or more portions of the operating phases of the engine of the aircraft to which the gas turbine engine is designed to be attached within a given mission.

[0335] In this regard, ground idle may refer to an engine operating phase in which the aircraft is stationary and in contact with the ground, but in which there are requirements for the engine to be operated. During idle, the engine may produce between 3% and 9% of the available thrust of the engine. In another non-limiting example, the engine may produce between 5% and 8% of the available thrust. In another non-limiting example, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to an engine operating phase in which the aircraft is propelled along the ground by the thrust produced by the engine. During taxiing, the engine may produce between 5% and 15% of the available thrust. In another non-limiting example, the engine may produce between 6% and 12% of the available thrust. In another non-limiting example, the engine may produce between 7% and 10% of the available thrust. Takeoff may refer to an engine operating phase in which the aircraft is propelled by the thrust produced by the engine. During the initial phase of the takeoff phase, the aircraft may be propelled while in contact with the ground. During a later phase of the takeoff phase, the aircraft may be propelled while not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In another non-limiting example, the engine may produce between 95% and 100% of the available thrust. In another non-limiting example, the engine may produce 100% of the available thrust.

[0336] Climb may refer to an engine operating phase in which the aircraft is propelled by the thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In another non-limiting example, the engine may produce between 80% and 95% of the available thrust. In another non-limiting example, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to the operating phase between takeoff and reaching cruise conditions within the aircraft flight cycle, reaching cruise conditions thus defining the start of the cruise phase or a part thereof of the aircraft flight. Additionally or alternatively, climb may refer to one or more nominal periods at a nominal point or during the aircraft flight cycle between takeoff and landing, in which a relative increase in altitude is required, which may require additional thrust requirements of the engine.

[0337] As used herein, the cruise conditions that may define the flight cruise phase (or a portion thereof) of an aircraft have their conventional meaning and will be readily understood by a person skilled in the art. In some examples, for a given gas turbine engine of an aircraft, the cruise conditions may refer to the operating point at which the engine cruises in the middle of a given mission (referred to in the industry as an "economic mission"), and the gas turbine engine is designed to be attached to the aircraft. In this regard, the mid-cruise can be regarded as a key point in the flight cycle of the aircraft, at which point 50% of the total fuel burned between the highest point of ascent and the start of descent has been burned (which may approximate the midpoint between the highest point of ascent and the start of descent in terms of time and / or distance). Thus, taking into account the number of engines provided for the aircraft, the cruise conditions can define the operating point, phase or a portion thereof of the flight that provides the thrust to ensure the steady-state operation of the aircraft (i.e., maintaining a constant altitude and / or a constant Mach number) or at least substantially steady-state operation (i.e., maintaining at least substantially a constant altitude and / or at least substantially a constant Mach number). For example, if the engine is designed to be attached to an aircraft with two engines of the same type, under cruise conditions, the engine can provide half of the total thrust required for the steady-state operation or at least substantially steady-state operation of the aircraft during mid-cruise.

[0338] In other words, for a given gas turbine engine of an aircraft, the cruise conditions can be defined as the operating point of the engine that provides a specified thrust under mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude), which is required to provide the steady-state operation, or at least substantially steady-state operation, of the aircraft to which the gas turbine engine is designed to be attached, in combination with any other engines on the aircraft, at a given mid-cruise Mach number. For any given gas turbine engine of an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine under cruise conditions can be clearly defined.

[0339] By way of example only, the forward speed under cruise conditions can be any point within the range from 0.7 Mach to 0.9 Mach, such as 0.75 to 0.85, such as 0.76 to 0.84, such as 0.77 to 0.83, such as 0.78 to 0.82, such as 0.79 to 0.81, such as approximately 0.8 Mach, approximately 0.85 Mach or within the range of 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, such as below 0.7 Mach or above 0.9 Mach.

[0340] By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere ISA) at altitudes within the following ranges: 10,000 m to 15,000 m, for example within the range of 10,000 m to 12,000 m, for example within the range of 10,400 m to 11,600 m (about 38,000 feet), for example within the range of 10,500 m to 11,500 m, for example within the range of 10,600 m to 11,400 m, for example within the range of 10,700 m (about 35,000 feet) to 11,300 m, for example within the range of 10,800 m to 11,200 m, for example within the range of 10,900 m to 11,100 m, for example approximately 11,000 m. The cruise conditions may correspond to the standard atmospheric conditions at any given altitude within these ranges.

[0341] By way of example only, the 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 ft (10,668 m). Under such cruise conditions, the engine may provide a known required net thrust level. Of course, the known required net thrust level depends on the engine and its intended application and may be a value within the range of, for example, 20 kN to 40 kN.

[0342] By way of a further example only, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 ft (11,582 m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level, of course, depends on the engine and its intended application and may be a value within the range of, for example, 35 kN to 65 kN.

[0343] In use, the gas turbine engine described and / or claimed herein may operate under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be installed to provide propulsive thrust.

[0344] In addition, those skilled in the art will immediately recognize that either or both of landing and approach refer to the operational phases within the flight cycle of an aircraft between cruise and landing, where approach in particular forms part of the landing and take-off (LTO) phases. During either or both of landing and approach, the engine can produce between 0% and 50% of available thrust. In additional non-limiting examples, the engine can produce between 25% and 40% of available thrust. In additional non-limiting examples, the engine can produce between 30% and 35% of available thrust. Additionally or alternatively, landing can refer to a nominal point within the flight cycle of an aircraft between take-off and landing where a relative reduction in altitude is required, and this can require a reduced thrust demand from the engine.

[0345] According to one aspect, there is provided an aircraft that includes a gas turbine engine as described herein and / or claimed. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise conditions according to this aspect can correspond to the operating point, phase, or a part thereof of the aircraft flight, as defined elsewhere herein.

[0346] According to one aspect, there is provided a method of operating a gas turbine engine as described herein and / or claimed. The operation can be carried out under any cruise conditions (e.g., in terms of thrust, atmospheric conditions, and Mach number) as defined elsewhere herein.

[0347] According to one aspect, there is provided a method of operating an aircraft that includes a gas turbine engine as described herein and / or claimed. The operation according to this aspect can include (or can be) an operation under any suitable conditions (e.g., at intermediate cruise of the aircraft), as defined elsewhere herein.

[0348] Those skilled in the art will understand that features or parameters described with respect to any one of the above aspects can be applied to any other aspect unless mutually exclusive. Additionally, unless mutually exclusive, any feature or parameter included or described herein can be applied to any aspect and / or combined with any other feature or parameter included or described herein.

[0349] Any parameter or value included or described herein may be applied to and / or combined with any one or more other parameters and / or values included or described herein, unless mutually exclusive. For example, a first parameter or value (e.g., parameter A) included or described herein may be applied to and / or combined with any one or more other parameters and / or values (e.g., parameter B; parameter C; and parameter D, etc.) included or described herein to express the product of their relationships. For example, those skilled in the art will understand that in the case where parameter A and parameter B are disclosed separately, the product of their relationships may be expressed, as needed, as, for example, A / B, B / A, B*A, or any such other application, combination, or function of parameter A relative to parameter B. BRIEF DESCRIPTION OF THE DRAWINGS

[0350] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0351] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0352] Figure 2 is a close-up cross-sectional side view of an upstream portion of a geared gas turbine engine;

[0353] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0354] Figure 4 is a close-up cross-sectional side view of an upstream portion of a direct drive gas turbine engine;

[0355] Figure 5 is a schematic representation of an example fuel system including a fuel return line;

[0356] Figure 6 is a schematic representation of an example fuel system including a fuel recirculation line;

[0357] Figure 7 is a schematic representation of another example fuel system including a fuel return line;

[0358] Figures 8 to 13 illustrates an example method of operating a gas turbine engine;

[0359] Figure 14 is a schematic representation of an example fuel system including a fuel recirculation line;

[0360] Figure 15 is a schematic representation of an example fuel system including a fuel recirculation line and a fuel return line;

[0361] Figure 16 is a schematic representation of an example fuel system including a fuel recirculation line; and

[0362] Figure 17 is a representation of an aircraft having a propulsion system that includes two gas turbine engines. Detailed Description

[0363] Figure 1 illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0364] In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where the compressed air is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as a burner 16, where the terms "combustion equipment 16" and "burner 16" may be used interchangeably herein. The resulting hot combustion products then expand through the high-pressure and low-pressure turbines 17, 19 before being discharged through the nozzle 20, thereby driving the high-pressure and low-pressure turbines to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 is generally used to apply an increased pressure to the bypass airflow B flowing through the bypass duct 22 such that the bypass airflow B is discharged through the bypass exhaust nozzle 18 to generally provide most of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0365] Figure 2 shows an exemplary arrangement of a geared fan gas turbine engine 10. The low-pressure turbine 19 (see Figure 1) The drive shaft 26 is coupled to the sun gear or sunwheel 28 of the epicyclic gear arrangement 30. Radially outward of the sun gear 28 and meshing therewith are a plurality of planet gears 32 which are connected together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28 while each planet gear 32 rotates about its own axis. The planet carrier 34 is coupled via a link 36 to the fan 23 so as to drive the fan to rotate about the engine axis 9. Radially outward of the planet gears 32 and meshing therewith is a ring gear or annulus 38 which is coupled via a link 40 to the fixed support structure 24.

[0366] Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may respectively denote the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the fan 23), and / or the turbine stage and the compressor stage connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be termed "intermediate pressure turbine" and "intermediate pressure compressor". In the case of using such alternative nomenclature, the fan 23 may be referred to as the first or lowest pressure compression stage.

[0367] In Figure 3 the epicyclic gearbox 30 is shown in more detail by way of example. Each of the sun gear 28, the planet gears 32 and the annulus 38 includes teeth around its periphery for meshing with other gears. However, for clarity, Figure 3 only an exemplary portion of the teeth is illustrated. Four planet gears 32 are illustrated, but it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of planetary epicyclic gearboxes 30 typically include at least three planet gears 32.

[0368] In Figure 2 and Figure 3 the epicyclic gearbox 30 illustrated by way of example is planetary, where the planet carrier 34 is coupled via a link 36 to the output shaft and the annulus 38 is fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of a further example, the epicyclic gearbox 30 may be a stellar arrangement where the planet carrier 34 remains fixed allowing the annulus (or ring gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the annulus 38. By way of another alternative example, the gearbox 30 may be a differential gearbox where both the annulus 38 and the planet carrier 34 are allowed to rotate.

[0369] It should be understood that, Figure 2 and Figure 3The arrangements shown are by way of example only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to locate the gearbox 30 within the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of a further example, the connecting members (such as Figure 2 the linkages 36, 40 in the example) between the gearbox 30 and other components of the engine (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of a further example, any suitable arrangement of bearings between the rotating and fixed components of the engine (for example, between the input and output shafts from the gearbox and the fixed structure such as the gearbox housing) may be used, and the present disclosure is not limited to Figure 2 the exemplary arrangement. For example, in the case where the gearbox 30 has a stellar arrangement (as described above), those skilled in the art will readily appreciate that the arrangement of the output linkages and the support linkages and the bearing positions will generally be different from Figure 2 the arrangement shown by way of example in

[0370] Accordingly, the present disclosure extends to gas turbine engines having any arrangement in the type of gearbox (such as stellar or planetary gears), the support structure, the input and output shaft arrangements, and the bearing positions.

[0371] Optionally, the gearbox may drive additional and / or alternative components (for example, an intermediate pressure compressor and / or a booster compressor).

[0372] 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 interconnecting shafts. By way of a further example, Figure 1 the gas turbine engine shown in

[0373] has split nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially external to the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to an engine 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 of the nozzles (whether mixed or split) may have a fixed or variable area. Figure 4 A cross-section of one such engine is shown in

[0374] Referring to Figure 4, A gas turbine engine is generally designated by 10 and has a main rotational axis 9. The engine 10 includes, in an axial flow series arrangement, an air intake 12, a propulsive fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines the air intake 12 and the exhaust nozzle 20.

[0375] In use, air entering the air intake 12 is accelerated by the fan 23 to produce two airflows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate pressure compressor 14, and the bypass airflow B passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the airflow A before delivering the air to the high pressure compressor 15 where further compression occurs.

[0376] The compressed air discharged from the high pressure compressor 15 is directed to the combustion equipment 16 where the compressed air is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as a burner 16, and the terms "combustion equipment 16" and "burner 16" may be used interchangeably herein. Then, the resulting hot combustion products expand through the high pressure turbine 17, the intermediate pressure turbine 19a, and the low pressure turbine 19 before being discharged through the nozzle 20, thereby driving the high pressure turbine, the intermediate pressure turbine, and the low pressure turbine 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 suitable interconnecting shafts.

[0377] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Additionally, the engine may include a gearbox disposed in the drive train from the turbine to the compressor and / or the fan.

[0378] Although the examples described relate to turbofan engines, the present disclosure may be applied to, for example, any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0379] The geometry of the gas turbine engine 10 and its components is defined by a conventional axis system, including an axial direction (aligned with the rotational axis 9), a radial direction (the direction from bottom to top in Figure 1 ), and a circumferential direction (perpendicular to the Figure 1 plane of the view). The axial direction, the radial direction, and the circumferential direction are perpendicular to each other.

[0380] The fuel F provided to the combustion apparatus 16 may include a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may include molecules from one or more of the chemical families of normal paraffins, isoparaffins, naphthenes, and aromatics. Additionally or alternatively, when blended with, mixed with, or replaced by an alternative fuel, the fuel F may contain renewable hydrocarbons produced from biological or non-biological resources, also referred to as sustainable aviation fuel (SAF). In each of the provided embodiments, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0381] Those skilled in the art understand SAF to refer to, for example, biofuels, renewable aviation fuels, renewable jet fuels, alternative fuels, or biojet fuels produced from biological or non-biological resources. SAF is understood to generally be synthesized from carbon-containing gases extracted from the atmosphere and / or captured from industrial processes; or from a wide range of sustainable feedstocks such as, for example, waste oils and fats; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata, and pennycress; abiogenic alternative fuels; jatropha; halophytes, and algae, rather than from fossil-based hydrocarbons. SAF is understood not to include fossil fuels.

[0382] The functional performance of a given fuel composition or fuel F blend for a given task may be defined, at least in part, by the ability of the Brayton cycle of the fuel-serviced gas turbine engine 10. Parameters that define the functional performance may include, for example, specific energy; energy density; thermal stability; and emissions, including gases and / or particulate matter. In this regard, particulate emissions may include soot particles produced by the combustion of the fuel F, also referred to as non-volatile particulate matter (nvPM). Any reference herein to soot or smoke equally applies to other types of particulate emissions known in the art. Gas 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 (VOC) produced by the combustion of the fuel F. Any reference herein to gas emissions equally applies to other types of gas emissions known in the art.

[0383] Relatively high specific energy (i.e., energy per unit mass) expressed in MJ / kg can at least partially reduce the takeoff weight and thus potentially provide a relative improvement in fuel efficiency. Relatively high energy density (i.e., energy per unit volume) expressed in MJ / L can at least partially reduce the takeoff fuel volume, which may be particularly important for volume-constrained missions or military operations involving fuel replenishment. Relatively high thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors and thus potentially provide a relative improvement in combustion efficiency. Reduced emissions (including particulate matter) can allow reduced contrail formation while reducing the environmental impact of a given mission. Other properties of the fuel may also be critical to functional performance. For example, a relatively low freezing point (°C) can allow long-term missions to optimize the flight profile; a minimum aromatic concentration (%) can ensure sufficient swelling of certain materials used to construct O-rings and seals that have previously been exposed to fuels with high aromatic content; and a maximum surface tension (mN / m) can ensure sufficient fuel spray breakup and atomization.

[0384] 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 composition or fuel blend. Fuels with a higher ratio of hydrogen atoms to carbon atoms can have higher specific energy in the absence of bond strain. For example, fossil-based hydrocarbon fuels can include molecules having from about 7 to 18 carbons, where a significant portion of a given composition is derived from molecules having from 9 to 15 carbons and having an average of 12 carbons.

[0385] Multiple sustainable aviation fuel blends have been approved for use. For example, some approved blends contain a blend ratio of up to 10% sustainable aviation fuel, while other approved blends contain a blend ratio of 10% to 50% sustainable aviation fuel (the remainder containing one or more fossil-based hydrocarbon fuels such as kerosene), with additional compositions pending approval. However, sustainable aviation fuel blends that are expected to include up to (and including) 100% sustainable aviation fuel (SAF) in the aviation industry will ultimately be approved for use. Any reference to "SAF" herein can be a fuel that is 100% SAF or a fuel that contains SAF, e.g., is a SAF blend.

[0386] Sustainable aviation fuel may include one or more of normal alkanes, isoalkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of the following: syngas; lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuel may include either or both of lower aromatic and sulfur content (relative to fossil-based hydrocarbon fuels). Additionally or alternatively, sustainable aviation fuel may include either or both of higher isoalkane and cycloalkane content (relative to fossil-based hydrocarbon fuels). In some examples, sustainable aviation fuel may include either or both of a density between 90% and 98% of the density of kerosene and a calorific value between 101% and 105% of the calorific value of kerosene.

[0387] In some examples, the aromatic and / or other non-alkane content of the sustainable aviation fuel, or blend, provided to the combustion apparatus 16 is relatively lower than that of kerosene. Sustainable aviation fuel may include an aromatic content, such as 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; such as 4%, 3%, 2%, 1%, or less than 1%; such as 0.75%, 0.5%, 0.25%, or less than 0.25%; such as 0.2%, 0.1%, or less than 0.1%; such as 0.01%, 0.001%, or 0%. Depending on one or more of preference, fuel feedstock or supplier, and variations in its components, the aromatic content of sustainable aviation fuel may be within a numerical value or range including the end values defined by or bounded by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as 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%.

[0388] At least in part due to the molecular structure of sustainable aviation fuel, sustainable aviation fuel may provide benefits including, for example, one or more of the following: higher specific energy (although, 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; lower NOx; and lower CO2 emissions, relative to fossil-based hydrocarbon fuels (e.g., when burned in the combustion apparatus 16). Thus, relative to fossil-based hydrocarbon fuels such as kerosene, sustainable aviation fuel may result in either or both of a relative reduction in fuel consumption and a relative reduction in maintenance costs.

[0389] An example fuel system 1000 including a fuel flow path from the fuel tank 50 to the burner 16 of the gas turbine engine 10 of the aircraft 1 is in Figure 5is schematically shown. In the example shown, engine 10 includes a gearbox 30 as shown in Figures 1 to 3 (however, it can be used in combination with other gas turbine engines, including those with a direct drive architecture). The fuel system 1000 includes a fuel supply system (including a fuel tank 50 and a pump 1002) that supplies fuel to the engines of the aircraft and a fuel management system 1500 that operates within the engine. The fuel management system 1500 manages the fuel temperature as well as the fuel flow rate, guiding the fuel through one or more heat exchangers 1004, 1006 of the engine's heat exchange system.

[0390] In the described embodiment, each engine 10 has its own fuel management system 1500. In other embodiments, a single fuel management system 1500 can manage the fuel supply to multiple engines and can include, for example, Figure 5 replicas of various elements for other engines as shown.

[0391] Fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low-pressure fuel supply pump 1002. The fuel then flows through a secondary fuel-oil heat exchanger 1004 and a primary fuel-oil heat exchanger 1006 before passing through an engine fuel pump 1003. The engine fuel pump 1003 can be described as a main fuel pump or a high-pressure fuel pump. The primary fuel-oil heat exchanger 1006 can be described as the main 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 can be described as a generator fuel-oil heat exchanger because the oil flowing through it can be used to cool and / or lubricate the generator of the engine 10, which is configured to provide electrical power supply for the aircraft (e.g., an integrated drive generator (IDG) of the engine 10). The illustrated fuel management system 1500 is arranged such that the fuel reaches the secondary fuel-oil heat exchanger 1004 before reaching the primary fuel-oil heat exchanger 1006.

[0392] The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured so that an oil flow is conveyed through each heat exchanger in addition to the fuel flow flowing therethrough. The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured so that heat can be transferred between the oil flowing therethrough and the fuel. Under cruising conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006, and the average temperature of the oil flow entering the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1004. In this way, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are each configured to transfer thermal energy from the oil flow to the fuel flow flowing therethrough under cruising conditions. The fuel system 1000 includes an electronic controller configured to control the operation of the heat exchangers 1004, 1006.

[0393] The primary fuel-oil heat exchanger 1006 may have oil passing therethrough for lubricating and / or cooling the main gearbox 30 of the gas turbine engine 10, and thus may be described as a main fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1004 may have oil passing therethrough for lubricating and / or cooling one or more components of the generator of the engine 10, and thus may be described as a generator fuel-oil heat exchanger.

[0394] The two oil streams may be physically separated from each other and optionally chemically different and / or have different flow rates. Thus, the oil flowing through the primary fuel-oil heat exchanger 1006 may be different from the oil flowing through the secondary fuel-oil heat exchanger 1004 .

[0395] Generally, at least a majority of the fuel that passes through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006, although either or each heat exchanger 1004, 1006 may be provided with a bypass to allow some of the fuel to avoid passing through the respective heat exchanger, e.g., Figure 5 The form of bypass conduit 1005 is shown. A valve (not shown) may determine the proportion of fuel passing through heat exchanger 1004 and the proportion of fuel passing through bypass conduit 1005.

[0396] The two heat exchangers 1004, 1006 can thus be described as being in series with each other with respect to the fuel flow. Generally speaking, at least most of the fuel passing through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006, although each heat exchanger 1004, 1006 can be provided with a bypass to allow some of the fuel in the fuel to avoid passing through the corresponding heat exchanger, for example in the form of a bypass conduit 1005. A valve can determine the proportion of fuel passing through the heat exchanger 1004 and the proportion of fuel passing through the bypass conduit 1005. Additionally or alternatively, one or more bypass conduits can also be provided to allow the oil to bypass one or more of the heat exchangers.

[0397] The secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006 are configured such that an oil flow is also conveyed through each heat exchanger. In the described embodiment, the oil flowing through one fuel-oil heat exchanger is different from the oil flowing through the other fuel-oil heat exchanger, but it should be understood that in other embodiments, the same oil can flow through one fuel-oil heat exchanger and then through the other fuel-oil heat exchanger. In such an embodiment, one or more bypass conduits for the oil or fuel can again be provided to allow the fluid to bypass one or more of the heat exchangers.

[0398] Thus, in the described embodiment, the two heat exchangers 1004, 1006 are in separate closed-loop systems with respect to the oil flow, i.e., the oil flowing through the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger is fluidly separated and can be chemically different from each other.

[0399] The fuel system 1000 further includes a modulating valve 1010 that is located downstream of the primary fuel-oil heat exchanger 1006 and is arranged to transfer or direct at least a portion of the fuel that has left the primary fuel-oil heat exchanger via a fuel return line 1011 to the fuel tank 50. The modulating valve 1010 is configured to modulate the fuel flow along the fuel return line 1011. The modulating valve 1010 can determine the proportion of fuel returned to the fuel tank 50 and the proportion of fuel that continues to the burner 16. It should be understood that the modulating valve 1010 can be located at substantially any suitable position downstream of at least one of the heat exchangers 1004, 1006. For example, the modulating valve 1010 can alternatively be located downstream of the secondary heat exchanger 1004 and upstream of the primary heat exchanger 1006. It should also be understood that the fuel system 1000 can alternatively include only a single fuel-oil heat exchanger, or can include more than two fuel-oil heat exchangers, and the modulating valve 1010 can be located downstream of at least one of the heat exchangers (either directly downstream or with one or more intermediate components of the fuel system 1000 between the modulating valve 1010 and the heat exchanger).

[0400] Figure 5 The heat exchangers shown and described herein are merely examples. The primary heat exchanger and / or the secondary heat exchanger may provide cooling for other aircraft systems in addition to the systems described above.

[0401] In the illustrated example, the modulating valve 1010 is located downstream of the primary fuel-oil heat exchanger 1006. In the illustrated example, the modulating valve 1010 is located upstream of the engine fuel pump 1003, but this is not required. The modulating valve 1010 is arranged to allow a controlled amount of fuel to return to the fuel tank 50.

[0402] The fuel system 1000 includes a plurality of temperature sensors (represented by a circular symbol around the capital letter T) configured to measure the temperature of the fuel at different locations in the fuel system 1000. In the illustrated example, the fuel system 1000 includes a temperature sensor located in the fuel tank 50 to measure the temperature of the fuel in the fuel tank 50. Additionally or alternatively, the fuel system 1000 may include a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1004, 1006, which can be used to determine or represent the temperature of the fuel in the fuel tank 50.

[0403] In the illustrated example, the fuel system 1000 includes two temperature sensors located downstream of the heat exchangers 1004, 1006. One of the temperature sensors is located in the fuel return line 1011, while the other temperature sensor is located downstream of the modulating valve 1010 and upstream of the burner 16. Either of the two temperature sensors can be used to determine or represent the temperature of the fuel delivered to the burner 16 and / or the temperature of the fuel returned to the fuel tank 50. It should be understood that only one of the temperature sensors located downstream of the heat exchangers 1004, 1006 may be provided. Depending on the position of the modulating valve 1010 relative to one or more heat exchangers (e.g., the primary heat exchanger 1006 and / or the secondary heat exchanger 1004), it should be understood that the temperature sensor can be disposed at any suitable location downstream of one or more heat exchangers to measure temperature data related to the temperature of the fuel delivered to the burner 16 and / or the temperature of the fuel returned to the fuel tank 50 (or to provide temperature data representative of that temperature).

[0404] Returning fuel to the fuel tank 50 provides a mechanism for controlling the fuel flow rate of the fuel system 1000, for example, to manage the heat load in the engine 10 and / or to control the fuel temperature at different locations in the fuel system 1000 (such as in the fuel tank 50 or at the inlet of the burner 16).

[0405] The modulation valve 1010 can 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., open and close) of the modulation valve 1010.

[0406] In one example, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of the temperature of the fuel in the fuel tank 50 (in Kelvin) to the temperature of the fuel delivered to the burner 16 (in Kelvin) 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 burner 16 may also be known or referred to as the ΔT ratio (or simply ΔT). The fuel temperature measured downstream of the heat exchangers 1004, 1006 (e.g., by a temperature sensor located downstream of one or both of the heat exchangers 1004, 1006) can be used to determine or represent the fuel temperature delivered to the burner 16 (i.e., the fuel temperature at the burner inlet). The temperature of the fuel in the fuel tank 50 can 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 modulation valve 1010 at least in part based on temperature data from at least one of the temperature sensors.

[0407] 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 burner 16 in Kelvin (K), where 0 °C is equal to 273.15 K.

[0408] The temperature of the fuel in the fuel tank 50 is typically maintained between -54 °C (219.15 K) and 65 °C (338.15 K). The lower limit of -54 °C is typically used to prevent the fuel from freezing in the fuel tank 50. Alternatively, the upper limit of the temperature of the fuel in the fuel tank 50 can be 55 °C (328.15 K), or 5 °C (278.15 K), or 0 °C (273.15 K). In some other examples, the upper limit of the temperature of the fuel in the fuel tank 50 can be 100 °C (373.15 K).

[0409] The maximum temperature of the fuel in the fuel tank can be set by the fuel flash point to reduce the generation of fuel vapor, which may ignite if there is an ignition source. When the fuel does not return to the fuel tank, from the perspective of engine operation, it may be beneficial for the fuel in the tank to be colder, but not so cold as to cause fuel icing problems in the engine fuel system. If the fuel tank temperature is low, it is possible to absorb more heat from the heat generated within the engine back into the more efficient fuel. The inventors have recognized that when using fuel return to the tank, the system boundary extends from just the engine to the combined engine and aircraft fuel tank, so a colder tank is not necessarily more beneficial, and the fuel in the tank can be maintained at a hotter temperature.

[0410] For sustainable aviation fuel (i.e., fuel that is 100% SAF or a SAF blend), the temperature of the fuel delivered to the burner 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 (such as 200 °C or 473.15 K). More preferably, the temperature of the fuel delivered to the burner 16 under cruise conditions can typically be at least 140 °C (413.15 K), or typically between 140 °C (413.15 K) and 180 °C (453.15 K) or higher (such as 200 °C or 473.15 K).

[0411] Alternatively, the temperature of the fuel delivered to the burner 16 under cruise conditions can be at least 120 °C, 125 °C, 130 °C, 135 °C, 140 °C, 145 °C, 150 °C, 155 °C, 160 °C, 165 °C, 170 °C, 175 °C, 180 °C, 185 °C, 190 °C, 195 °C or 200 °C, or a fuel temperature defined between any two of these values.

[0412] The temperature of "the fuel delivered to the burner", the temperature of "the fuel supplied to the burner", and the temperature of "the fuel at the burner inlet" can be used interchangeably herein and should be considered to refer to the same temperature.

[0413] The temperature of the fuel delivered to the burner 16 under cruise conditions can be defined as the average value under steady-state cruise conditions for at least 5 minutes and optionally for 10 minutes. These average temperatures do not include transient spikes in temperature, which can be defined as fluctuations in the fuel temperature during operation, typically an increase in temperature. Each fluctuation can last no 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 control the temperature of the fuel delivered to the burner 16.

[0414] The modulation 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 a temperature of less than -54 °C / 120 °C. The inventors have recognized that this range can be achieved by using SAF, where the fuel temperature at the burner inlet is not limited to 120 °C. This may be otherwise the case for fossil-kerosene fuel due to the risk of fuel thermal decomposition. In this example, the minimum temperature of the fuel in the fuel tank is limited to -54 °C to avoid the risk of freezing. In this example (and in other similar examples where only an upper limit is given), the value of ΔT will be greater than zero.

[0415] Generally, heating the fuel to a higher temperature when delivered to the burner can provide improved engine thermodynamic efficiency. The inventors have determined that using SAF allows for a higher fuel temperature at the burner inlet without the risk of significant fuel thermal decomposition, which may otherwise occur if fossil kerosene fuel is used. The inventors have determined how to advantageously control the fuel / heat management system of the engine to take advantage of this difference between SAF and kerosene. Raising the temperature of the fuel also helps to allow more heat to be transferred out of the oil within the heat exchanger for improved cooling.

[0416] In some examples, the modulation 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 a temperature of less than -54 °C / 140 °C. This corresponds to a minimum temperature of the fuel supplied to the burner of 140 °C and a minimum temperature of the fuel in the fuel tank of -54 °C. The inventors have determined that it is possible to operate using SAF in this state as it can be advantageously heated to a higher temperature to improve thermodynamic efficiency without the risk of such thermal decomposition.

[0417] In some examples, the modulation 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 temperature between -54 °C / 180 °C and -54 °C / 120 °C. This corresponds to a fuel temperature at the burner inlet between 120 °C and 180 °C while the minimum temperature of the fuel in the fuel tank remains at -54 °C. The maximum limit of 180 °C may be beneficial to reduce the risk of fuel decomposition that may occur even when using SAF.

[0418] In some examples, the modulation 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 temperature between -54°C / 180°C and -54°C / 140°C. This corresponds to the fuel temperature at the burner inlet being between 140°C and 180°C while the minimum temperature of the fuel in the fuel tank is maintained at -54°C. This again makes better use of the thermal properties of the SAF to improve the thermodynamic efficiency while avoiding the risk of fuel decomposition at too high a temperature.

[0419] In some examples, the modulation 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 between -54°C / 180°C and -65°C / 65°C. The lower limit corresponds to the maximum temperature of 180°C for the fuel at the burner inlet as described above, 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 in this example is not configured to operate under conditions where heat is removed from the fuel, i.e., the fuel is not colder at the burner inlet compared to the fuel temperature in the fuel tank.

[0420] In some examples, the modulation 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 between -54°C / 180°C and 100°C / 140°C. Compared to the example in the previous paragraph, the lower upper limit of 0.90 corresponds to the maximum temperature of the fuel in the fuel tank being 100°C as described above, where the minimum temperature of the fuel at the burner inlet is 140°C, to make use of the greater thermal properties of the SAF.

[0421] In some examples, the modulation 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 between -54°C / 180°C and 65°C / 140°C. The upper limit of 0.82 corresponds to the maximum temperature of the fuel in the fuel tank being 65°C as described above, where the minimum temperature of the fuel at the burner inlet is 140°C, to make use of the greater thermal properties of the SAF to improve the thermodynamic efficiency.

[0422] In some examples, the modulation 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 between -54°C / 140°C and 65°C / 65°C. The upper limit of 1.00 corresponds to the maximum value at which the fuel in the fuel tank is not hotter than the fuel in the burner. The lower limit corresponds to the temperature of the fuel at the burner inlet being 140°C or higher to utilize the thermal properties of SAF to improve the thermodynamic efficiency, while the minimum temperature at the fuel tank is -54°C to reduce the risk of freezing.

[0423] In some examples, the modulation 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 between -54°C / 140°C and 65°C / 140°C. This corresponds to the same lower limit as in the example of the previous paragraph. The upper limit of 0.82 corresponds to the maximum temperature of the fuel in the fuel tank being 65°C as described above, where the minimum temperature of the fuel at the burner inlet is 140°C to utilize the greater thermal properties of SAF to increase the thermodynamic efficiency.

[0424] In some examples, the modulation 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 between -54°C / 140°C and 65°C / 180°C. This corresponds to the same lower limit as in the example of the previous paragraph. The upper limit of 0.75 corresponds to the maximum temperature of the fuel in the fuel tank being 65°C as described above, and the value of the maximum temperature of the fuel at the burner inlet being 180°C as described above. The maximum value at the burner inlet is only used in some examples when the fuel temperature in the fuel tank is relatively high to avoid a large temperature difference between them.

[0425] Figure 8 A method 2000 of operating a gas turbine engine 10 is shown. The method 2000 includes modulating 2001 the fuel flow along the fuel return line 1011 using the modulation valve 1010 such that the ratio of the temperature of the fuel in the fuel tank to the temperature of the fuel delivered to the burner (ΔT) is defined as in any of the examples described above with respect to the fuel system 1000 or as defined elsewhere herein.

[0426] In another example, the modulation 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 the ratio of ΔT during engine 10 cruise to ΔT during startup is less than 0.56. The electronic controller of the fuel system 1000 is configured to control the operation of the modulation valve 1010 accordingly to achieve this value or any other value of ΔT during cruise to ΔT during startup as defined herein.

[0427] Operating the modulation valve such that the ratio of ΔT during cruise to ΔT at engine 10 startup is less than 0.56 corresponds to a value less than ((-54 °C / 120 °C) / 1.00). Thus, it corresponds to a value achievable using a fuel temperature of 120 °C or higher at the burner inlet, which can be achieved by leveraging the improved thermal properties of SAF compared to fossil fuels. The value 0.56 corresponds to the minimum value of ΔT set by the minimum fuel temperature of -54 °C in the fuel tank to reduce the risk of freezing when reducing the fuel temperature at the burner inlet to at least 120 °C. The maximum value of ΔT at startup is 1, which corresponds to the maximum value achievable when the fuel in the fuel tank is not hotter than the fuel in the burner.

[0428] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during cruise of engine 10 to ΔT at startup is less than 0.53 (i.e., (-54 °C / 140 °C) / 1.00). This corresponds to the value in the example of the previous paragraph, except that the fuel temperature at the burner inlet is at least 140 °C to further leverage the thermal properties of SAF to improve the thermodynamic efficiency.

[0429] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during cruise of engine 10 to ΔT at startup is between 0.48 and 0.56. This corresponds to a temperature between ((-54 °C / 180 °C) / 1.00 and ((-54 °C / 120 °C) / 1.00). The upper limit of 0.56 corresponds to the upper limit discussed above. The lower limit of 0.48 corresponds to the maximum temperature of the fuel at the burner inlet as described above of 180 °C to reduce the risk of fuel thermal decomposition, where the minimum temperature of the fuel in the fuel tank is -54 °C to reduce the risk of freezing.

[0430] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during cruise of engine 10 to ΔT at startup is between 0.48 and 0.53. This corresponds to a temperature between (-54 °C / 180 °C) / 1.00 and (-(-54 °C / 140 °C) / 1.00). The upper limit of 0.53 corresponds to the upper limit of further leveraging the thermal properties of SAF discussed above. The lower limit of 0.48 corresponds to the lower limit discussed in the previous paragraph.

[0431] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.88. This corresponds to temperatures between ((-54 °C / 180 °C) / (100 °C / 100 °C)) and ((100 °C / 120 °C) / (-54 °C / 160 °C)). The inventors have determined that the bounds of the ratio of ΔT during cruise to ΔT during startup can be determined by the following expression:

[0432] (((min_tank_cruise / max_combustor_cruise) / (max_tank_startup / min_combustor_startup @ max tank startup)) / ((max_tank_cruise / min_combustor_cruise) / (min_tank_startup / max_combustor_startup)))

[0433] The lower bound of 0.48 corresponds to a minimum temperature of -54 °C for the fuel in the tank to reduce the risk of freezing, and a maximum temperature of 180 °C for the fuel at the burner inlet during cruise to reduce the risk of fuel thermal decomposition and improve thermodynamic efficiency. The maximum value of ΔT during startup is 1 because the heat exchanger is not arranged to cool the fuel before it reaches the burner. Thus, the value of ΔT during startup is limited by the minimum temperature of the fuel at the burner inlet, which is not less than the corresponding maximum value of the fuel in the tank (e.g., 100 °C) when calculating the maximum value of ΔT. The upper bound of 1.88 corresponds to a maximum temperature of 100 °C for the fuel in the tank, a minimum temperature of 120 °C for the fuel at the burner inlet to utilize the thermal properties of the SAF, a minimum temperature of -54 °C for the fuel in the tank to reduce the risk of freezing, and a maximum temperature of 160 °C for the fuel at the burner inlet during startup.

[0434] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.71. This corresponds to temperatures between ((-54 °C / 180 °C) / (65 °C / 65 °C)) and ((65 °C / 120 °C) / (-54 °C / 160 °C)). These bounds correspond to the bounds in the previous paragraph, but with a maximum temperature of 65 °C for the fuel in the fuel tank.

[0435] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.70. This corresponds to temperatures between ((-54 °C / 180 °C) / (100 °C / 100 °C)) and ((100 °C / 120 °C) / (-54 °C / 120 °C)). In this example, ΔT during startup can be ΔT during startup when the aircraft is on the ground. The bounds in this paragraph correspond to the bounds in the previous paragraph, except that the maximum temperature of the fuel in the fuel tank is 100 °C and the maximum temperature of the fuel at the burner inlet during startup is 120 °C.

[0436] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.55. This corresponds to temperatures between ((-54 °C / 180 °C) / (65 °C / 65 °C)) and ((65 °C / 120 °C) / (-54 °C / 120 °C)). In this example, ΔT during startup can be ΔT during startup when the aircraft is on the ground. These bounds are consistent with the bounds in the previous paragraph, except that the maximum temperature of the fuel in the tank is 65 °C.

[0437] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.88. This corresponds to temperatures between ((-54 °C / 180 °C) / (100 °C / 100 °C)) and ((100 °C / 120 °C) / (-54 °C / 160 °C)). In this example, ΔT during startup can be ΔT during startup when the aircraft is in the air. These bounds are consistent with the bounds 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 at the burner inlet during startup is 160 °C.

[0438] In some examples, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 to the fuel tank 50 such that the ratio of ΔT during engine 10 cruise to ΔT during startup is between 0.48 and 1.71. This corresponds to temperatures between ((-54 °C / 180 °C) / (65 °C / 65 °C)) and ((65 °C / 120 °C) / (-54 °C / 160 °C)). In this example, ΔT during startup can be ΔT during startup when the aircraft is in the air. These bounds are consistent with the bounds in the previous paragraph, except that the maximum temperature of the fuel in the tank is 65 °C.

[0439] In any of the examples described herein, the fuel temperature at the burner inlet during cruise can be from 120°C to 180°C, and preferably from 140°C to 180°C. In any of the examples described herein: the fuel temperature at the burner inlet during startup can be from -40°C to 160°C. When starting on the ground, the fuel temperature at the burner inlet can be from -40°C to 120°C. When starting in the air, the fuel temperature at the burner inlet can be from 0°C to 160°C.

[0440] The temperature ranges in the previous paragraph may correspond to a ground start, where the engine starts all components at ambient temperature. For an in-air (i.e., in-flight) start, the temperature ranges in the previous paragraph correspond to a re-ignition situation. In this case, the engine is running and then restarted. This may involve descending the aircraft from cruise altitude to the windmill re-ignition envelope. During this time, the temperature of the engine core components is decreasing, but there may still be some residual heat in the engine when re-ignition will occur. Thus, the temperature ranges for a ground restart and an in-air restart are different.

[0441] Figure 9 A method 2100 of operating a gas turbine engine 10 is shown. Method 2100 includes modulating 2101 the fuel flow along a fuel return line 1011 using a modulating valve 1010 such that the ratio of ΔT during cruise of the engine 10 to ΔT during startup is as defined in any of the examples above or as defined elsewhere herein.

[0442] In another example, the modulating valve 1010 is additionally or alternatively arranged to initiate returning fuel to the fuel tank 50 when the fuel delivered to the burner 16 (or having passed through one or both of the heat exchangers 1004, 1006) is at a temperature of at least 120°C. When the fuel temperature is below 120°C, the modulating valve 1010 can prevent fuel from returning to the fuel tank 50. An electronic controller of the fuel system 1000 is configured to control the operation of the modulating valve 1010 accordingly. In other examples, the modulating valve can be arranged to initiate returning fuel to the fuel tank 50 when the fuel delivered to the burner 16 (or having passed through one or both of the heat exchangers 1004, 1006) is at a temperature of at least 140°C. This can further utilize the thermal properties of SAF. In additional examples, the modulating valve can be arranged to initiate returning fuel to the fuel tank when the fuel that has passed through one or both of the heat exchangers is at a temperature between 120°C and 180°C, and preferably between 140°C and 180°C.

[0443] The temperature of the fuel is measured downstream of one or both of the heat exchangers 1004, 1006. In this example, a temperature sensor in the fuel return line 1011 is used to measure the temperature of the fuel, but it should be understood that the temperature of the fuel can be measured using a temperature sensor located at substantially any location downstream of one or both of the heat exchangers 1004, 1006. The temperature of the fuel that has passed through one or both of the heat exchangers 1004, 1006 can be used to determine or represent 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 modulation valve 1010 at least in part based on temperature data from a temperature sensor located downstream of one or both of the heat exchangers 1004, 1006.

[0444] In this example, the modulation valve 1010 is arranged to initiate the return of the fuel to the fuel tank 50 when the fuel that has passed through one or both of the heat exchangers 1004, 1006 is at a temperature of at least 120 °C (or at least 140 °C; or between 120 °C and 180 °C; or between 140 °C and 180 °C) and provides an indication of the operating conditions, but this is not necessary.

[0445] The operating conditions can be one or more operating conditions, including but not limited to the proportion of sustainable aviation fuel (SAF) in the fuel, the thermal stability of the fuel, the coking level of the fuel, the oxygen content of the fuel, and the sulfur content of the fuel.

[0446] The operating conditions can be detected on the aircraft, for example during flight, for example using one or more sensors and / or other measurements. Data from the sensors can be provided to the electronic controller of the fuel system 1000 to control the operation of the modulation valve 1010. Alternatively, the operating conditions can be looked up, manually entered, or transmitted to the aircraft 1 (for example to the electronic engine controller of the fuel system 1000) before or during flight.

[0447] For example, trace substances or species that are naturally present in the fuel or added as tracers can be used to determine fuel properties, such as the percentage of sustainable aviation fuel in the fuel or whether the fuel is kerosene. The measurement 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 coking level of the fuel, for example by measuring the accumulation of deposits on the upper surface of the piezoelectric crystal, which will cause a change in the vibration mode. Other fuel properties, such as sulfur content, naphthalene content, aromatic content, and hydrogen-carbon ratio, can be determined by measuring substances present in the exhaust gas emitted by the gas turbine engine 10 during use (for example, during flight).

[0448] In this example, the modulation valve 1010 is also arranged to modulate the fuel flow along the fuel return line 1011 at least partially based on the amount of fuel remaining in the fuel tank, but this is not necessary.

[0449] The modulation valve 1010 is additionally or alternatively arranged to modulate the fuel flow along the fuel return line 1011 such that, after the fuel returns to the fuel tank 50, the equilibrium temperature of the fuel in the fuel tank 50 does not exceed the 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 modulation valve 1010 can remain open and / or how much fuel can return to the fuel tank 50 (and optionally further modulate the mass flow rate of the fuel returning to the fuel tank 50 using the modulation valve 1010). A logical example that can be used to operate the modulation valve 1010 in this manner is provided below:

[0450] (M 箱 +M FRTT )·C p (T eq )·T eq =M 箱 ·C p (T 箱 )·T 箱 +M FRTT ·C p (T FRTT )·T FRTT where M 箱 is the mass of the fuel in the fuel tank 50, M FRTT is the mass of the fuel returning to the fuel tank 50 along the fuel return line 1011, T eq is the equilibrium temperature of the fuel mixture in the fuel tank 50 after the fuel returns to the fuel tank 50, T 箱 is the current temperature of the fuel in the fuel tank 50 before mixing, T FRTT is the temperature of the fuel returning to the fuel tank 50 along the fuel return line 1011, and C p is the specific heat capacity of the fuel at the respective temperatures (e.g., T eq , T 箱 , T FRTT ). The mass of the fuel to be returned to the fuel tank 50 can be determined based on the desired equilibrium temperature and the current temperatures of the fuel in the fuel tank 50 and the fuel returning 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 the fuel to be returned to the fuel tank 50 is determined, the modulation valve 1010 can be operated accordingly to return the required mass of fuel to the fuel tank 50.

[0451] The temperature of the fuel that has passed through one or both of the heat exchangers 1004, 1006 can be used to determine or represent the temperature of the fuel returning to the fuel tank 50. In this example, a temperature sensor in the fuel return line 1011 is used to measure the temperature of the fuel returning to the fuel tank 50, but it should be understood that a temperature sensor located at substantially any location downstream of one or both of the heat exchangers 1004, 1006 can be used to measure the temperature of the fuel returning to the fuel tank 50. 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 can be used to measure the temperature of the fuel in the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulation valve 1010 at least in part based on temperature data from at least one of the temperature sensors.

[0452] It should be understood that any suitable alternative logic (e.g., at least in part based on the amount of fuel remaining in the fuel tank 50) can be used to control the operation of the modulation valve 1010.

[0453] Figure 10 A method 2200 of operating a gas turbine engine 10 is shown. The method 2200 includes initiating 2201 the return of fuel to the fuel tank using the modulation valve 1010 when the fuel that has passed through the heat exchanger is at a temperature of at least 120°C or within any of the other ranges described above with respect to the fuel system 1000 or elsewhere herein.

[0454] In another example, the modulation valve 1010 is additionally or alternatively arranged to modulate the fuel flow along the fuel return line 1011 such that the ratio of the fuel mass returning to the tank 50 to the fuel mass delivered to the burner 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 modulation valve 1010. The modulation valve 1010 can be arranged to modulate the fuel flow along the fuel return line 1011 such that the ratio of the fuel mass returning to the tank 50 to the fuel mass delivered to the burner 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.

[0455] SAF generally has a higher calorific value than conventional fuels such as kerosene or kerosene-based fuels. For example, compared to the typical calorific value of 43.2 MJ / kg for a kerosene-based fuel such as Jet-A, the calorific value of SAF may be between 43.5 MJ / kg and 44 MJ / kg. Therefore, a smaller mass flow rate of SAF is required to provide the same amount of fuel power input to the burner as provided by using a conventional fuel.

[0456] Additionally, SAF generally has higher thermal stability than conventional fuels and can thus operate at higher temperatures before undergoing fuel decomposition (due to thermal degradation). Therefore, SAF is able to absorb more heat than conventional fuels before undergoing fuel decomposition.

[0457] Returning fuel to the fuel tank 50 typically occurs when additional fuel is used for purposes other than combustion, such as managing the heat load in the engine 10. For example, the fuel can be used to cool the oil in the engine 10 via heat exchangers 1004, 1006, and / or drive one or more actuators for more generally operating the engine 10 or other components in the aircraft 1. Since SAF can absorb more heat than conventional fuels and has a higher calorific value, the inventors have determined that different ratios of the mass of fuel returned to the fuel tank to the mass of fuel delivered to the burner can provide improved performance. For example, by using SAF, the increased calorific value of the fuel may result in less mass of fuel being burned, thus allowing a greater percentage of the fuel to be returned to the tank compared to using conventional fossil fuels.

[0458] Figure 11 A method 2300 of operating a gas turbine engine 10 is shown. The method 2300 includes modulating 2301 the fuel flow along the fuel return line 1011 such that the ratio of the mass of fuel returned to the tank 50 to the mass of fuel delivered to the burner 16 is between 0 and 9 under cruise conditions, as described above with respect to the fuel system 1000. The ratio of the mass of fuel returned to the tank 50 to the mass of fuel delivered to the burner 16 can be within any of the other ranges defined above or elsewhere herein.

[0459] In another example, the modulating valve 1010 is additionally or alternatively arranged to prevent fuel having a temperature of 180 °C or higher from returning to the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulating valve 1010 accordingly.

[0460] The temperature of the fuel is measured downstream of one or both of the heat exchangers 1004, 1006. In this example, a temperature sensor in the fuel return line 1011 is used to measure the temperature of the fuel, but it should be understood that a temperature sensor located at substantially any position downstream of one or both of the heat exchangers 1004, 1006 can be used to measure the temperature of the fuel. The temperature of the fuel that has passed through one or both of the heat exchangers 1004, 1006 can be used to determine or represent 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 modulating valve at least in part based on temperature data from a temperature sensor located downstream of one or both of the heat exchangers 1004, 1006.

[0461] In this example, the modulation valve 1010 is also arranged to prevent fuel from returning to the fuel tank when the temperature of the fuel in the fuel tank 50 is at a predetermined upper threshold temperature. As described above, the temperature of the fuel in the fuel tank 50 is typically maintained between -54 °C (219.15 K) and 100 °C (373.15 K). More preferably, it can be maintained between -54 °C (219.15 K) and 65 °C (338.15 K). The upper limit of the temperature of the fuel in the fuel tank 50 can be any suitable value between the temperature pairs in the previous two sentences. For example, the upper limit can be 55 °C (328.15 K), or 5 °C (278.15 K), or 0 °C (273.15 K). In some examples, the upper limit of the fuel in the fuel tank can 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.

[0462] When the fuel in the fuel tank 50 is at a predetermined upper threshold temperature, by preventing the fuel from returning to the fuel tank 50, the threshold temperature will not be exceeded due to the fuel that has passed through one or both of the heat exchangers 1004, 1006 returning to the fuel tank 50. 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 can be used to measure the temperature of the fuel in the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulation valve 1010 at least in part based on the temperature data from at least one of the temperature sensor providing data indicating the temperature of the fuel returning to the fuel tank 50 and the temperature sensor providing data indicating the temperature of the fuel in the fuel tank 50.

[0463] In this example, the modulation valve 1010 is also arranged to modulate the fuel flow along the fuel return line 1011 at least in part based on the amount of fuel remaining in the fuel tank, but this is not necessary.

[0464] The modulation valve 1010 can be arranged to enable fuel to return to the fuel tank 50 if the temperature of the fuel returning to the fuel tank 50 is lower than 180 °C and the temperature of the fuel in the fuel tank 50 is lower than a predetermined upper threshold temperature. If both of these conditions are met, the modulation valve 1010 is arranged to modulate the fuel flow along the fuel return line 1011 such that after the fuel returns to the fuel tank 50, the equilibrium temperature of the fuel in the fuel tank 50 does not exceed the predetermined upper threshold temperature. The electronic controller of the fuel system 1000 can determine how long the modulation valve 1010 can remain open and / or how much fuel can return to the fuel tank 50 (and optionally further modulate the mass flow rate of the fuel returning to the fuel tank 50 using the modulation valve 1010). A logical example that can be used to operate the modulation valve 1010 in this way is provided below:

[0465] (M tank +M FRTT )·C p (T eq )·T eq =M tank ·C p (T tank )·T tank +M FRTT ·C p (T FRTT )·T FRTT Where M tank is the mass of the fuel in the fuel tank 50, M FRTT is the mass of the fuel returning to the fuel tank 50 along the fuel return line 1011, T eq is the equilibrium temperature of the fuel mixture in the fuel tank 50 after the fuel has returned to the fuel tank 50, T tank is the current temperature of the fuel in the fuel tank 50 before mixing, T FRTT is the current temperature of the fuel returning to the fuel tank 50 along the fuel return line 1011, and C p is the specific heat capacity of the fuel at the respective temperatures (e.g., T eq , T tank , T FRTT ). The mass of the fuel to be returned to the fuel tank 50 can be determined based on the desired equilibrium temperature and the current temperatures of the fuel in the fuel tank 50 and the fuel returning to the fuel tank 50 along the fuel return line 1011, ensuring that the equilibrium temperature does not exceed the predetermined upper threshold temperature of the fuel in the fuel tank 50. Once the mass of the fuel to be returned to the fuel tank 50 has been determined, the modulation valve 1010 can be operated accordingly to return the required mass of fuel to the fuel tank 50.

[0466] The temperature of the fuel that has passed through one or two heat exchangers 1004, 1006 can be used to determine or represent the temperature of the fuel returned to the fuel tank 50. In this example, a temperature sensor in the fuel return line 1011 is used to measure the temperature of the fuel returned to the fuel tank 50, but it should be understood that a temperature sensor located at substantially any location downstream of one or two heat exchangers 1004, 1006 can be used to measure the temperature of the fuel returned to the fuel tank 50. 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 can be used to measure the temperature of the fuel in the fuel tank 50. The electronic controller of the fuel system 1000 is configured to control the operation of the modulation valve 1010 at least in part based on temperature data from at least one of the temperature sensors.

[0467] It should be understood that any suitable alternative logic for controlling the operation of the modulation valve 1010 (e.g., at least in part based on the amount of fuel remaining in the fuel tank 50) can be used.

[0468] Figure 12 A method 2400 of operating a gas turbine engine 10 is shown. The method 2400 includes preventing 2401 fuel having a temperature of 180 °C or higher from returning to the fuel tank 50 using the modulation valve 1010, as described above with respect to the fuel system 1000. The method 2400 can include preventing 2401 fuel having a temperature within any of the ranges defined above or elsewhere herein from returning to the fuel tank 50. The method can include preventing fuel from returning if the fuel will exceed the upper limit of the tank temperature defined above or as defined elsewhere herein.

[0469] The fuel system 1500 can be configured to perform any one or more of the methods 2000, 2100, 2200, 2300, 2400. If the fuel system is equipped with a suitable fuel recirculation line, it can also perform the Figure 13 methods described below.

[0470] Figure 6 Another example fuel system 1100 is schematically shown, which includes a fuel flow path from the fuel tank 50 to other components of the fuel system 1100. The fuel system 1100 is substantially similar to the fuel system 1000 described above with reference to Figure 5 wherein like reference numerals denote like elements.

[0471] The fuel system 1100 includes a fuel recirculation line 1111. The fuel recirculation line 1111 is arranged to recirculate fuel from a first point on the fuel flow path to a second point on the fuel flow path. In this example, the fuel flow path may be referred to as the main fuel flow path as it extends from the fuel tank to the burner. The second point on the fuel flow path is upstream of the first point. The first point corresponds to the inlet of the fuel recirculation line 1111. The second point corresponds to the outlet of the fuel recirculation line 1111. The fuel system 1100 includes a modulating valve 1110 arranged to modulate the fuel flow along the fuel recirculation line 1111 between the inlet and the outlet of the fuel recirculation line 1111. The fuel system 1100 includes a temperature sensor located in the fuel recirculation line 1111, but this is not necessary. The fuel system 1100 may also include a fuel return line (not shown for clarity), as described above with respect to Figure 5 the fuel system 1000 shown, but this is not necessary.

[0472] In the example shown, the first point or inlet of the fuel recirculation line 1111 is downstream of the primary fuel - oil heat exchanger 1106 on the fuel flow path. The first point is downstream of the fuel pump 1103. The modulating valve 1110 is also on the fuel flow path, but it should be understood that the modulating valve 1110 could alternatively be located along the fuel recirculation line 1111. The second point or outlet of the fuel recirculation line 1111 is upstream of the primary fuel - oil heat exchanger 1106.

[0473] The inlet and outlet of the fuel recirculation line 1111 can alternatively be located at any suitable location on the fuel flow path. For example, the inlet of the fuel recirculation line 1111 can be upstream of the fuel pump 1003, or downstream of the secondary fuel - oil heat exchanger 1104, or downstream of the fuel pump 1102. The outlet of the fuel recirculation line 1111 can be upstream of the fuel pump 1103, or upstream of the secondary fuel - oil heat exchanger 1104, or upstream of the fuel pump 1102. When the outlet of the fuel recirculation line 1111 is upstream of the fuel pump 1102, the inlet of the recirculation line can be upstream of the pump 1103. It should also be understood that the fuel system 1100 can alternatively include only a single fuel - oil heat exchanger, or can include more than two fuel - oil heat exchangers, and the first or second point of the fuel recirculation line 1111 can be located at any suitable location relative to one or more of the heat exchangers (upstream or downstream, directly or indirectly having one or more intervening components on the main fuel flow path). A fuel recirculation line 1111 can be provided for any given fuel - oil heat exchanger.

[0474] 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 (i.e., to recirculate some of the fuel along at least a portion of the fuel flow path). In the example shown, the modulating valve 1110 is arranged to modulate the fuel flow along the fuel recirculation line 1111 such that a certain amount of fuel passes through the primary fuel-oil heat exchanger 1106 multiple times before being delivered to the burner 16. This can allow additional heat transfer to the same fuel and increase the temperature of the fuel before it is delivered to the burner 16. The temperature of the fuel can be increased to at least 120 °C before the fuel is delivered to the burner 16. Thus, recirculating the fuel through the fuel recirculation line 1111 using the modulating valve 1110 can allow hotter fuel to be delivered to the burner 16, which can improve the efficiency of the engine 10. Additionally, the amount (e.g., mass) of fuel used to manage the heat load in the engine 10 (e.g., via the heat exchangers 1104, 1106) can be reduced, which in turn can reduce (or eliminate, depending on operating conditions) the amount of fuel that needs to be returned to the fuel tank 50 when managing the heat load in the engine 10. Reducing or eliminating the amount of fuel returned to the fuel tank 50 can also keep the fuel in the fuel tank 50 cooler before it enters the fuel flow path, which can improve the cooling of the oil in the fuel-oil heat exchangers 1104, 1106.

[0475] The inlet of the modulating valve 1110 and / or the fuel recirculation pipe 1111 can be located downstream of the fuel pump 1103 such that the recirculated fuel passes through the primary fuel-oil heat exchanger 1106 and the fuel pump 1103 multiple times rather than just through the primary heat exchanger 1106. For a given shaft speed of the engine 10, recirculating the fuel through the fuel pump 1103 can allow for more adjustable control of the fuel flow to the burner 16, noting that the pump speed (or a limited set of pump speed options) is typically set by the shaft speed (e.g., the pump speed can be proportionally related to the shaft speed). For example, if the fuel flow rate required to maintain the current aircraft thrust is lower than the pump speed for the required shaft speed, a larger proportion of the fuel can be recirculated through the pump such that the flow through the pump is higher than the flow to the burner 16. Without changing the pump speed, by increasing the recirculation amount, it is also possible to provide a lower flow rate of fuel to the burner 16 for a higher calorific value fuel. Usually, no fuel is wasted because the fuel is always recirculated and burned in the burner.

[0476] In the example shown, the modulating valve 1110 is arranged to actively control the amount of fuel recirculated along the fuel recirculation line 1111. The electronic controller of the fuel system 1000 is configured to control the operation of the modulating valve 1110. The active control can be performed based on one or more parameters, such as:

[0477] · Core shaft speed and engine thrust demand;

[0478] ·The current fuel temperature at one or more locations (e.g., downstream of one or more heat exchangers along a fuel flow path and / or along a fuel recirculation line);

[0479] ·The calorific value of the fuel;

[0480] ·The fan speed;

[0481] ·The fuel flow rate to the burner (commonly referred to as WFE, i.e., the weight of the main engine fuel flow);

[0482] ·The fuel viscosity; and

[0483] ·The main or engine fuel pump speed, or speed option.

[0484] Additionally or alternatively, the speed of the fuel pump 1103 can be adjusted, or the fuel flow rate can be increased (and thus the heat transfer per unit volume of fuel through the heat exchangers 1104, 1106 can be reduced) or the fuel flow rate can be decreased (and thus the heat transfer per unit volume of fuel through the heat exchangers 1104, 1106 can be increased).

[0485] Alternatively, the control of the amount of fuel recirculated along the fuel recirculation line 1111 using the modulation valve 1110 can be deactivated. Instead, a set or fixed proportion of the fuel in the main fuel flow path can be recirculated along the fuel recirculation line 1111 via the modulation valve 1110. For example, the ratio of the mass of recirculated fuel to the mass of fuel delivered to the burner can be between 0 and 9. More preferably, the ratio of the mass of recirculated fuel to the mass of fuel delivered to the burner under cruise conditions can be greater than 4 and less than or equal to 9, and even more preferably greater than 4 and less than or equal to 4.9 under cruise conditions. Thus, it can be the same as the fuel ratio returned to the tank as defined above, such that anything described herein applicable to the fuel returned to the fuel tank can also be applicable to this example, in which the fuel is recirculated along the fuel recirculation line 1111.

[0486] In some examples, the fuel recirculation line 1111 is arranged to supply fuel to one or more additional aircraft and / or engine mechanisms (not shown) before rejoining the fuel flow path. One or more additional aircraft mechanisms can thus be located on the fuel recirculation line 1111. The additional aircraft mechanisms can include nacelle anti-icing systems, actuators, turbine casing cooling systems, or any other suitable aircraft mechanism.

[0487] Figure 13 A method 2500 of operating a gas turbine engine 10 is shown. 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.

[0488] Figure 7 Another example fuel system 1200 is schematically shown, which includes a fuel flow path from a fuel tank 50 to other components of the fuel system 1200. In the example shown, the engine 10 is a direct drive engine as Figure 4 shown. However, Figure 7 the fuel system 1200 of Figure 6 can be used with other engine architectures, such as a geared architecture. The fuel system 1200 includes a fuel supply system (including a fuel tank 50 and a pump 1202) that supplies fuel to the engine 10 of the aircraft 1 and a fuel management system 1600 that operates within the engine 10. In the described embodiment, each engine 10 has its own fuel management system 1600. In other embodiments, a single fuel management system 1600 can manage the fuel supply to multiple engines and can include, for example,

[0489] replicas of various elements for other engines as

[0490] shown. Figure 7 Fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low pressure fuel supply pump 1202. The fuel then flows through a primary fuel-oil heat exchanger 1204 before flowing through an engine fuel pump 1203. The engine fuel pump 1203 can be described as a main fuel pump or a high pressure fuel pump. Then, at least a portion of the fuel flows through a secondary fuel-oil heat exchanger 1206 via a line 1207 that branches off from the main fuel flow path (between the fuel tank 50 and the burner 16), and at least a portion of the fuel flows to the burner 16 without flowing through the secondary fuel-oil heat exchanger 1206. The primary fuel-oil heat exchanger 1204 can be described as a main fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1206 can be described as a servo fuel-oil heat exchanger. The fuel management system 1600 is arranged such that the fuel reaches the primary fuel-oil heat exchanger 1204 before reaching the secondary fuel-oil heat exchanger 1206.

[0491] The primary fuel - oil heat exchanger 1204 and the secondary fuel - oil heat exchanger 1206 are configured such that an oil flow is also conveyed through the heat exchangers. The primary fuel - oil heat exchanger 1204 and the secondary fuel - oil heat exchanger 1206 are configured such that heat can be transferred between the oil flowing therethrough and the fuel. Under cruise conditions, the average temperature of the oil flow entering the primary fuel - oil heat exchanger 1204 and the secondary fuel - oil heat exchanger 1206 is respectively higher than the average temperature of the fuel entering the primary fuel - oil heat exchanger 1204 and the secondary fuel - oil heat exchanger 1206. 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 the oil flow to the fuel flow passing therethrough under cruise conditions. The fuel system 1200 includes an electronic controller configured to control the operation of the heat exchangers 1204, 1206.

[0492] In various arrangements of the fuel management system 1600, the oil flows through the secondary fuel - oil heat exchanger 1206 before flowing through the primary fuel - oil heat exchanger 1204 and does not flow through any engine components that would increase the temperature between them. Thus, the oil is hotter when entering the secondary fuel - oil heat exchanger 1206 than when entering the primary fuel - oil heat exchanger 1204. In contrast, the fuel flows through the primary fuel - oil heat exchanger 1204 before flowing through the secondary fuel - oil heat exchanger 1206. In this arrangement, the fuel temperature leaving the secondary fuel - oil heat exchanger 1206 is higher than the fuel temperature leaving the primary fuel - oil heat exchanger 1204. The engine fuel pump 1203 can alternatively be placed upstream of the secondary fuel - oil heat exchanger 1206 (or upstream where the line 1207 rejoins the main fuel flow path), which means it is not exposed to these additionally elevated fuel temperatures.

[0493] The fuel management system 1600 can be provided with a bypass to allow some of the fuel to avoid passing through the respective heat exchangers, for example, in the form of a bypass duct, the inlet of which is located upstream of the primary fuel - oil heat exchanger 1204 and the outlet of which is located downstream of the primary fuel - oil heat exchanger 1204 and upstream of the secondary fuel - oil heat exchanger 1206, as referred to above Figure 5 described. A valve can determine the proportion of the fuel passing through the heat exchanger 1204 and the proportion of the fuel passing through the bypass duct.

[0494] The fuel system 1200 also includes a modulating valve 1210 that is located downstream of the main fuel - oil heat exchanger 1204 and is arranged to divert or direct at least a portion of the fuel that has exited the primary fuel - oil heat exchanger 1204 via a fuel return line 1211 to the fuel tank 50. The modulating valve 1210 is configured to modulate the fuel flow along the fuel return line 1211. The modulating valve 1210 can determine the proportion of fuel returned to the fuel tank 50 and the proportion of fuel that continues to the burner 16. It should be understood that the modulating valve 1210 can be located at substantially any suitable position downstream of at least one of the heat exchangers 1204, 1206. For example, the modulating valve 1210 can alternatively be located upstream of the secondary heat exchanger 1206 (e.g., upstream of where the line 1207 rejoins the main fuel flow path). It should also be understood that the fuel system 1200 can alternatively include only a single fuel - oil heat exchanger, or can include more than two fuel - oil heat exchangers, and the modulating valve 1210 can be located downstream of at least one of the heat exchangers (either directly downstream or with one or more intermediate components of the fuel system 1200 between the modulating valve 1210 and the heat exchanger). The above - mentioned primary and secondary heat exchangers are merely an example and can provide cooling to any suitable engine system or component.

[0495] In the illustrated example, the modulating valve 1210 is located downstream of the primary fuel - oil heat exchanger 1204, the main fuel pump 1203, the inlet of the line 1207 leading to the secondary fuel - oil heat exchanger 1206, and the outlet of the line 1207 rejoining the main fuel flow path from the secondary fuel - oil heat exchanger 1206. It should be understood that the modulating valve 1210 can alternatively be located upstream of the fuel pump 1203.

[0496] The fuel system 1200 includes a plurality of temperature sensors (represented by a circular symbol around the capital letter T) that are configured to measure the temperature of the fuel at different locations in the fuel system 1200. In the illustrated example, the fuel system 1200 includes a temperature sensor located in the fuel tank 50 to measure the temperature of the fuel in the fuel tank 50. Additionally or alternatively, the fuel system 1200 can include a temperature sensor located downstream of the fuel tank 50 and upstream of the heat exchangers 1204, 1206, which can be used to determine or represent the temperature of the fuel in the fuel tank 50.

[0497] In the illustrated example, the fuel system 1200 includes two temperature sensors located downstream of the heat exchangers 1204, 1206. One of the temperature sensors is located in the fuel return line 1211, while the other temperature sensor is located downstream of the modulation valve 1210 and upstream of the burner 16. Either or both of the two temperature sensors can be used to determine or represent the temperature of the fuel delivered to the burner 16 and / or the temperature of the fuel returned to the fuel tank 50. It should be understood that only one of the temperature sensors located downstream of the heat exchangers 1204, 1206 can be provided. Depending on the position of the modulation valve 1210 relative to one or more heat exchangers (such as the primary heat exchanger 1204 and / or the secondary heat exchanger 1206), it should be understood that the temperature sensor can be disposed at any suitable position downstream of the one or more heat exchangers in order to measure temperature data related to the temperature of the fuel delivered to the burner 16 and / or the temperature of the fuel returned to the fuel tank 50 (or provide temperature data representative of that temperature).

[0498] Returning fuel to the fuel tank 50 provides a mechanism for controlling the fuel flow rate of the fuel system 1200, for example, to manage the heat load in the engine 10 and / or to control the fuel temperature at different locations in the fuel system 1200 (such as in the fuel tank 50 or at the burner 16 inlet).

[0499] The modulation valve 1210 can 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 Figure 5 the fuel system 1000 shown. The fuel system 1200 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulation valve 1210. It should also be understood that Figures 8 to 12 any one or more of the methods 2000, 2100, 2200, 2300, 2400 shown in Figure 7 and described above with respect to the fuel system 1000 can equally be used Figure 13 with the fuel system 1200 shown in

[0500] Figure 14 Another example fuel system 1300 is schematically shown, which includes a fuel flow path from the fuel tank 50 to other components of the fuel system 1300. The fuel system 1300 is substantially similar to the fuel system 1200 described above with reference to Figure 7 wherein like reference numerals represent like elements.

[0501] The fuel system 1300 includes a fuel recirculation line 1311. The fuel recirculation line 1311 is arranged to recirculate fuel from a first point on the fuel flow path to a second point on the fuel flow path. In this example, the fuel flow path is the main fuel flow path extending from the fuel tank to the burner. The second point on the fuel flow path is upstream of the first point. The first point corresponds to the inlet of the fuel recirculation line 1311. The second point corresponds to the outlet of the fuel recirculation line 1311. The fuel system 1300 includes a modulating valve 1310 arranged to modulate the fuel flow along the fuel recirculation line 1311 between the inlet and the outlet of the fuel recirculation line 1311. The fuel system 1300 includes a temperature sensor located in the fuel recirculation line 1311, but 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 Figure 7 the fuel system 1200 shown, but this is not required.

[0502] In the example shown, the first point or inlet of the fuel recirculation line 1311 is downstream of the primary fuel-oil heat exchanger 1304 on the fuel flow path. The first point is downstream of the fuel pump 1303. The first point is downstream of the inlet of the line 1307 leading to the secondary fuel-oil heat exchanger 1306 and downstream of the outlet of the line 1307 rejoining the main fuel flow path from the secondary fuel-oil heat exchanger 1306 (between the fuel tank 50 and the burner 16). Thus, the first point is downstream of the flow path into and out of the secondary heat exchanger 1306. The second point is upstream of the flow path into and out of the secondary heat exchanger. The modulating valve 1310 is also on the main fuel flow path, but it should be understood that the modulating valve 1310 could alternatively be located at another position on the fuel flow path or along the fuel recirculation line 1311. The second point or outlet of the fuel recirculation line 1311 is upstream of the primary fuel-oil heat exchanger 1304.

[0503] The inlet and outlet of the fuel recirculation line 1311 may alternatively be located at any suitable location on the fuel flow path. For example, the inlet of the fuel recirculation line 1311 may be located downstream of the primary heat exchanger 1304 but upstream of the fuel pump 1303, or downstream of the primary heat exchanger 1304 and the fuel pump 1303 but upstream of the inlet of the line 1307 leading to the secondary heat exchanger 1306, or downstream of the outlet of the line 1307 that rejoins the main fuel flow path from the secondary heat exchanger 1306. The outlet of the fuel recirculation line 1311 may be located upstream of the inlet of the line 1307 leading to the secondary heat exchanger 1306, or upstream of the fuel pump 1303, or upstream of the fuel pump 1302. It should also be understood that the fuel system 1300 may alternatively include only a single fuel-oil heat exchanger, or may include more than two fuel-oil heat exchangers, and the first or second point of the fuel recirculation line 1311 may be located at any suitable location relative to one or more heat exchangers (upstream or downstream, directly or indirectly having one or more intermediate components in the fuel flow path). A fuel recirculation line 1311 may be provided for any given fuel-oil heat exchanger.

[0504] The fuel recirculation line 1311 is arranged to allow some fuel to flow multiple times along at least a portion of the fuel flow path before being delivered to the burner 16 (i.e., to recirculate some of the fuel along at least a portion of the fuel flow path). In the example shown, the modulating valve 1310 is arranged to modulate the fuel flow along the fuel recirculation line 1311 such that a certain amount of fuel passes through the primary fuel-oil heat exchanger 1304 multiple times before being delivered to the burner 16. This can allow additional heat transfer to the same fuel and increase the temperature of the fuel before it is delivered to the burner 16. The temperature of the fuel can be raised to at least 120°C before the fuel is delivered to the burner 16. In some examples, the temperature of the fuel can be raised to at least 140°C, or within a range between 120°C and 180°C or between 140°C and 180°C, before the fuel is delivered to the burner 16. Thus, recirculating the fuel through the fuel recirculation line 1311 using the modulating valve 1310 can allow hotter fuel to be delivered to the burner 16, which can improve the efficiency of the engine 10. Additionally, the amount (e.g., mass) of fuel used to manage the heat load in the engine 10 (e.g., through the heat exchangers 1304, 1306) can be reduced, which in turn can reduce (or eliminate, depending on the operating conditions) the amount of fuel that needs to be returned to the fuel tank 50 when managing the heat load in the engine 10. Reducing or eliminating the amount of fuel returned to the fuel tank 50 can also keep the fuel in the fuel tank 50 cooler before it enters the fuel flow path, which can improve the cooling of the oil in the fuel-oil heat exchangers 1304, 1306.

[0505] In the example shown, the inlet of modulation valve 1310 and / or fuel recirculation pipe 1311 is also located downstream of fuel pump 1303 such that the recirculated fuel passes through both primary fuel - oil heat exchanger 1304 and fuel pump 1303 multiple times. For a given shaft speed of engine 10, recirculating fuel through fuel pump 1303 allows for more adjustable control of the fuel flow to burner 16, noting that the pump speed (or a limited set of pump speed options) is typically set by the shaft speed (e.g., the pump speed may be proportionally related to the shaft speed). For example, if the fuel flow required to maintain the current aircraft thrust is lower than the pump speed for the required shaft speed, a larger proportion of the fuel can be recirculated through the pump such that the flow through the pump is higher than the flow to burner 16. Without changing the pump speed, by increasing the recirculation amount, a lower flow of fuel can also be provided to burner 16 for higher calorific value fuels.

[0506] Modulation valve 1310 can be arranged to modulate the fuel flow along fuel recirculation line 1311 in any suitable manner, e.g., as described above with respect to Figure 7 fuel system 1200 shown. Fuel system 1300 includes an electronic controller configured to control the operation (e.g., open and close) of modulation valve 1310. It should also be understood that Figure 13 the method 2500 shown in Figure 14 and described above with respect to fuel system 1100 can likewise be performed using

[0507] fuel system 1300 shown.

[0508] 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. One or more additional aircraft mechanisms can thus be located on fuel recirculation line 1311. The additional aircraft mechanisms can include nacelle anti - ice systems, actuators, turbine housing cooling systems, or any other suitable aircraft mechanism.

[0508] In Figure 15 and Figure 16 additional example fuel systems 1400, 1700 are schematically shown which include a fuel flow path from fuel tank 50 to burner 16 of gas turbine engine 10 of aircraft 1. Fuel system 1400 is similar to fuel systems 1000, 1100 described above with reference to Figure 5 and Figure 6 wherein like reference numerals represent like elements. Fuel system 1700 is similar to fuel systems 1200, 1300 described above with reference to Figure 7 and Figure 14 wherein like reference numerals represent like elements.

[0509] In fuel systems 1400, 1700, the modulation valves 1410, 1710 are arranged to divert or direct at least a portion of the fuel that has left the fuel-oil heat exchanger to the fuel tank 50 and / or to recycle the fuel to an upstream location on the fuel flow path. Thus, it can perform Figures 8 to 13 the method of any one or more of the figures.

[0510] In Figure 15 the example shown, the modulation valve 1410 of the fuel system 1400 is located downstream of the secondary fuel-oil heat exchanger 1406 and the pump 1403. The modulation valve 1410 is arranged to divert or direct at least a portion of the fuel from the pump 1403 back to the fuel tank 50 via the fuel return line 1411b and is also arranged to recycle the fuel from the pump 1403 to a point upstream of the secondary heat exchanger 1406 via the fuel recirculation line 1411a. It should also be understood that the modulation valve 1410 can alternatively be located at any suitable location in the fuel system 1400, such as on the fuel flow path.

[0511] The modulation valve 1410 can be arranged to modulate the fuel flow along the fuel recirculation line 1411a and the fuel return line 1411b in any suitable manner, such as, for example, as described above with respect to Figure 5 , Figure 6 , Figure 7 and Figure 14 the fuel systems 1000, 1100, 1200, 1300 shown. The fuel system 1400 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulation valve 1410. It should also be understood that Figures 8 to 13 the methods 2000, 2100, 2200, 2300, 2400, 2500 shown in and described above with respect to the fuel systems 1000, 1100, 1200, 1300 can equally be performed using Figure 15 the fuel system 1400 shown.

[0512] In Figure 16In the example shown, the modulating valve 1710 of the fuel system 1700 is located downstream of the secondary fuel - oil heat exchanger 1706. The fuel leaving the secondary heat exchanger 1706 does not re - engage the main fuel flow path (between the fuel tank 50 and the burner 16). The modulating valve 1710 is arranged to divert or direct at least a portion of the fuel leaving the secondary heat exchanger 1706 back to the fuel tank 50 via the fuel return line 1711b, and is also arranged to recycle the fuel leaving the secondary heat exchanger 1706 to a point upstream of the primary fuel - oil heat exchanger 1704 on the main fuel flow path via the fuel recirculation line 1711a. It should also be understood that the modulating valve 1710 may alternatively be located at any suitable location, such as on the main fuel flow path, such that the fuel leaving the secondary heat exchanger 1706 can re - engage the main fuel flow path before encountering the modulating valve 1710.

[0513] The modulating valve 1710 can be arranged to modulate the fuel flow along the fuel recirculation line 1711a and the fuel return line 1711b in any suitable manner, e.g., as described above with respect to Figure 5 , Figure 6 , Figure 7 and Figure 14 the fuel systems 1000, 1100, 1200, 1300 shown. The fuel system 1700 includes an electronic controller configured to control the operation (e.g., opening and closing) of the modulating valve 1710. It should also be understood that Figures 8 to 13 the methods 2000, 2100, 2200, 2300, 2400, 2500 shown in Figure 16 and described above with respect to the fuel systems 1000, 1100, 1200, 1300 can equally be performed using

[0514] the fuel system 1700 shown in

[0515] Figure 17 An aircraft 1 is shown on which two gas turbine engines 10 of the present disclosure are mounted, one on each wing. Fuel F is supplied from a fuel tank 50 to the gas turbine engines 10. In this example, the fuel tank 50 includes a set of interconnected fuel tanks. In Figure 17In the example shown, the fuel tank consists of a primary fuel tank located in the aircraft fuselage and smaller fuel tanks located in each wing. In other examples, the aircraft 1 may have only a single fuel tank - it should be understood that many different fuel tank layouts can be envisioned, and the example depicted is provided for ease of description and is not intended to be limiting.

[0516] It should be understood that the present invention is not limited to the above examples, and various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature can be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and sub - combinations of one or more of the features described herein.

Claims

1. A gas turbine engine for an aircraft, comprising: Burner; a fuel-oil heat exchanger arranged to receive fuel and transfer heat from the oil to the fuel so as to raise the temperature of the fuel at the burner inlet to at least 120°C; a fuel recirculation line arranged to recirculate at least some of the fuel on a fuel flow path from a first point on the fuel flow path to a second point on the fuel flow path, the second point being upstream of the first point; and A modulator valve is arranged to modulate fuel flow along the fuel recirculation line.

2. The gas turbine engine according to claim 1, wherein: The fuel recirculation line is 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.

3. The gas turbine engine of claim 1 , further comprising a fuel pump located on the fuel flow path, and wherein: The fuel recirculation line is arranged to recirculate fuel from a first point downstream of the fuel pump to a second point upstream of the fuel pump.

4. The gas turbine engine according to claim 3, wherein: The fuel pump is located downstream of the fuel-oil heat exchanger on the fuel flow path.

5. The gas turbine engine according to claim 1, wherein: The fuel recirculation line is arranged to supply fuel to one or more additional aircraft and / or engine mechanisms.

6. The gas turbine engine according to claim 5, wherein: The one or more additional mechanisms include one or more of a nacelle anti-icing system, an actuator, a bleed valve, and a thermal management modulator valve.

7. The gas turbine engine according to claim 1, wherein: The fuel-oil heat exchanger is arranged to transfer heat from the oil to the fuel so as to raise the temperature of the fuel at the burner inlet to at least 140°C.

8. The gas turbine engine according to claim 1, wherein: The fuel-oil heat exchanger is arranged to transfer heat from the oil to the fuel so as to raise the fuel temperature at the burner inlet to between 120°C and 180°C, preferably to between 140°C and 180°C.

9. The gas turbine engine according to claim 1, wherein: The modulation valve is arranged to modulate the fuel flow along the fuel recirculation line such that a ratio of a recirculated fuel mass to a fuel mass delivered to the combustor is between 0 and 9 under cruise conditions.

10. The gas turbine engine of claim 1, wherein: The modulation valve is arranged to modulate the fuel flow along the fuel recirculation line so that the ratio of the recirculated fuel mass to the fuel mass delivered to the combustor is between 2.3 and 9 under cruise conditions, preferably greater than 4 and less than or equal to 9 under cruise conditions, or further preferably greater than 4 and less than or equal to 4.9 under cruise conditions.

11. A method of operating a gas turbine engine, the gas turbine engine comprising: Burner; a fuel-oil heat exchanger arranged to receive fuel and transfer heat from the oil to the fuel so as to raise the temperature of the fuel at the burner inlet to at least 120°C; a fuel recirculation line arranged to recirculate at least some of the fuel on a fuel flow path from a first point on the fuel flow path to a second point on the fuel flow path, the second point being upstream of the first point; and a modulator valve arranged to modulate the flow of fuel along the fuel recirculation line, Therein, the method includes modulating the fuel flow along the fuel recirculation line using the modulation valve.

12. The method according to claim 11, wherein: The method includes modulating the flow of fuel along the fuel recirculation line from a first point downstream of the fuel-oil heat exchanger to a second point upstream of the fuel-oil heat exchanger using the modulation valve.

13. The method according to claim 11, wherein: The gas turbine engine also includes a fuel pump located on the fuel flow path, and wherein the method includes modulating the fuel flow along the fuel recirculation line from a first point downstream of the fuel pump to a second point upstream of the fuel pump using the modulation valve.

14. The method according to claim 13, wherein: The fuel pump is located downstream of the fuel-oil heat exchanger on the fuel flow path.

15. The method according to claim 11, wherein: The method includes supplying fuel to one or more additional aircraft and / or engine mechanisms via the fuel recirculation line.

16. The method according to claim 15, wherein: The one or more additional mechanisms include one or more of a nacelle anti-icing system, an actuator, a bleed valve, and a thermal management modulator valve.

17. The method according to claim 11, wherein: The method comprises transferring heat from the oil to the fuel so as to increase the temperature of the fuel at the burner inlet to at least 140°C.

18. The method according to claim 11, wherein: The method comprises transferring heat from the oil to the fuel so as to raise the temperature of the fuel at the burner inlet to between 120°C and 180°C, preferably to between 140°C and 180°C.

19. The method according to claim 11, wherein: The method includes modulating the fuel flow along the fuel recirculation line using the modulation valve such that a ratio of a recirculated fuel mass to a fuel mass delivered to the combustor is between 0 and 9 under cruise conditions.

20. The method according to claim 11, wherein: The method includes modulating the fuel flow along the fuel recirculation line using the modulation valve so that a ratio of a recirculated fuel mass to a fuel mass delivered to the combustor is between 2.3 and 9 under cruise conditions, preferably greater than 4 and less than or equal to 9 under cruise conditions, or further preferably greater than 4 and less than or equal to 4.9 under cruise conditions.

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