Fuel hydraulic actuation

Abstract

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor. The engine comprises a fan located upstream of the engine core and arranged to be driven by the mandrel. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass conduit located radially outward of the engine core with a bypass ratio greater than or equal to 4 defined as a ratio of a mass flow rate of flow through the bypass conduit to a mass flow rate of flow through the core under cruise conditions. The engine includes a fuel supply system and a plurality of actuators. The fuel supply system is arranged to supply fuel for combustion in the combustor and to supply fuel to hydraulically drive at least ten of the plurality of actuators by fuel.

Description

[0001] Cross - reference to related applications

[0002] This specification is based on and claims the benefit of priority of United Kingdom Patent Application No. 2319159.6, filed on December 14, 2023, the entire content of which is incorporated herein by reference. Background of the Invention Field of the Invention

[0003] The present disclosure relates to an aircraft actuation system and to a method of controlling the actuation system in a fuel - hydraulic manner such that one or more actuators of the system are actuated in a fuel - hydraulic manner.

[0004] Description of Related Technologies

[0005] In the aviation industry, there is a trend towards using fuels different from the traditional kerosene - based jet fuels commonly used at present. These fuels may have different fuel properties relative to petroleum - based hydrocarbon fuels. Therefore, it is necessary to consider the fuel properties of these new fuels and to adjust the gas turbine engine itself and the method of operating the gas turbine engine. 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] An engine core including a turbine, a combustor, a compressor, and a shaft connecting the turbine to the compressor;

[0008] A fan located upstream of the engine core and arranged to be driven by the shaft, the fan including a plurality of fan blades;

[0009] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0010] A plurality of actuators; and

[0011] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the combustor and to supply fuel to drive at least ten of the plurality of actuators in a fuel - hydraulic manner.

[0012] The present inventors have recognized that using fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel characteristics, and these different fuel characteristics can enable actuators to be driven in a fuel-hydraulic manner. In particular, some fuels can be heated to higher temperatures and used to drive at least one more actuator than traditional fuels without significantly increasing the risk of fuel thermal degradation (e.g., fuel varnishing or fuel coking) within the actuator.

[0013] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0014] The fuel supply system can include a fuel return tank valve. A fuel return tank actuator can be used to control the fuel return tank valve. The fuel supply system can be arranged to supply fuel to drive the fuel return tank actuator in a fuel-hydraulic manner.

[0015] The fuel supply system can be arranged to supply fuel to drive at least eleven or at least twelve actuators out of a plurality of actuators in a fuel-hydraulic manner.

[0016] The engine can include a variable stator vane system. At least two actuators out of a plurality of actuators can be part of the variable stator vane system and can actuate / move the vanes in use. The fuel supply system can be arranged to supply fuel to drive at least two actuators out of a plurality of actuators that are part of the variable stator vane system in a fuel-hydraulic manner.

[0017] The engine can include a turbine casing cooling - TCC - system. At least two actuators out of a plurality of actuators can be part of the turbine casing cooling system. The fuel supply system can be arranged to supply fuel to drive at least two actuators out of a plurality of actuators that are part of the TCC system in a fuel-hydraulic manner.

[0018] The engine can include a ventilation valve actuator. The fuel supply system can be arranged to supply fuel to drive the ventilation valve actuator in a fuel-hydraulic manner.

[0019] The engine may include an engine thermal management system. The engine thermal management system may be arranged to cool the oil that is used to cool and lubricate one or more engine components including bearings and a gearbox (if present). The engine thermal management system may be arranged to control the fuel temperature entering the burner. The engine thermal management system may include a plurality of heat exchangers; for example, one or more air-oil heat exchangers and one or more fuel-oil heat exchangers. At least one of the plurality of actuators may be part of the engine thermal management system. The fuel supply system is arranged to supply fuel to hydraulically drive at least one of the plurality of actuators within the engine thermal management system. Various actuators within the engine thermal management system may be used to control the fluid flow rate through one or more of the heat exchangers within the engine thermal management system, such as the air, oil, or fuel flow rate.

[0020] The engine may include a generator thermal management system. At least one of the plurality of actuators is part of the generator thermal management system. The fuel supply system is arranged to supply fuel to hydraulically drive at least one of the plurality of actuators within the generator thermal management system.

[0021] The engine may include a hydraulic mechanical unit. At least three of the plurality of actuators may be part of the hydraulic mechanical unit. The fuel supply system may be arranged to supply fuel to hydraulically drive at least three of the plurality of actuators that are part of the hydraulic mechanical unit.

[0022] At least four of the plurality of actuators may be part of the hydraulic mechanical unit. The fuel supply system may be arranged to supply fuel to hydraulically drive at least four of the plurality of actuators that are part of the hydraulic mechanical unit.

[0023] The engine may include a bleed air valve. At least one of the plurality of actuators may be configured to actuate the bleed air valve. The fuel supply system may be arranged to supply fuel to hydraulically drive at least one of the plurality of actuators that is configured to actuate the bleed air valve.

[0024] Thus, at least ten fuel hydraulic actuators (i.e., actuators hydraulically driven by fuel) may include one or more of the following:

[0025] · At least one variable stator vane - VSV - actuator that is arranged to adjust the position of the stator vanes of the engine;

[0026] · At least one actuator that is arranged to actuate a valve within the turbine case cooling - TCC - system of the engine;

[0027] · At least one actuator that is arranged to actuate a valve within the hydraulic mechanical unit - HMU - of the engine;

[0028] · at least one actuator arranged to actuate a ventilation valve;

[0029] · at least one actuator arranged to actuate a nacelle anti - ice valve or any other anti - ice or de - ice valve;

[0030] · at least one actuator arranged to actuate a bleed air valve;

[0031] · at least one actuator arranged to actuate a valve of a regulating assembly;

[0032] · at least one actuator arranged to actuate a high - level valve for controlling the airflow from a compressor;

[0033] · at least one actuator arranged to actuate a valve for extracting bleed air from non - compressor components of the engine;

[0034] · at least one actuator arranged to actuate an auxiliary power unit - APU - bleed air valve (however, in some embodiments, the APU may be separated from the engine, and the APU actuator may thus not be classified as part of the engine. Fuel from a part of the fuel supply path upstream of the engine may be used to actuate such APU actuators. However, in other embodiments, the APU may be associated with the engine or be part of the engine, and in such cases the AP actuator may be classified as part of the engine);

[0035] · an engine start valve actuator;

[0036] · an isolation valve actuator;

[0037] · at least one actuator arranged to actuate a valve of an engine thermal management system; and / or

[0038] · at least one actuator arranged to actuate a valve of a generator thermal management system.

[0039] For example, an engine thermal management system typically includes a plurality of servomotors (also known as servo mechanisms), and a servo mechanism is a rotary or linear actuator that allows precise control of angular or linear position, speed, and / or acceleration in a mechanical system. In a given engine thermal management system, one, some, or all of the servo mechanisms present may be actuated in a fuel - hydraulic manner. One or more servo mechanisms may include sensors arranged to provide position feedback. A dedicated controller may be provided for the engine thermal management system to control one or more servo mechanisms. This also applies to the generator thermal management system.

[0040] The spool may directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct - drive turbofan engine.

[0041] The turbomachine may include a gearbox that receives an input from a spool and outputs drive to a fan to drive the fan at a rotational speed lower than the spool, such that the engine is a geared turbomachine.

[0042] The engine may include a plurality of actuated engine systems. The actuated engine systems may include one or more of the following: a thermal management system, a turbine case cooling system, a generator thermal management system, an engine thermal management system, a bleed air system, a variable stator vane - VSV - system, a nacelle environmental control system, an anti - ice system, and a bearing chamber ventilation system.

[0043] The fuel supply system may be arranged to supply fuel to hydraulically drive at least three of the plurality of actuated engine systems. The fuel supply system may be arranged to supply fuel to hydraulically drive at least four, five, six, or at least seven of the actuated engine systems.

[0044] It should be understood that each actuated engine system may include a plurality of actuators, and not all actuators of a given system may be fuel - hydraulic. Thus, hydraulically driving a given system includes actuating at least one actuator of the system, but not necessarily all actuators.

[0045] At least one of the plurality of actuated engine systems may include at least two of the plurality of actuators. The fuel supply system may be arranged to supply fuel to hydraulically drive at least one of the at least two actuators within at least one of the actuated engine systems. The fuel supply system may be arranged to supply fuel to hydraulically drive each of the at least two actuators within at least one of the actuated engine systems.

[0046] According to a second aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0047] an engine core including a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor;

[0048] a fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0049] a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0050] a plurality of actuators; and

[0051] Fuel supply system;

[0052] The method includes:

[0053] Using a fuel supply system to supply fuel for combustion in a burner; and

[0054] Using a fuel supply system to supply fuel to hydraulically drive at least ten of a plurality of actuators.

[0055] The method of the second aspect can be performed using the engine of the first aspect.

[0056] According to a third aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0057] An engine core including a turbine, a burner, a compressor, and a spool connecting the turbine to the compressor;

[0058] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0059] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0060] A plurality of actuated engine systems including a thermal management system and a turbine casing cooling system;

[0061] And

[0062] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in a burner and to supply fuel to hydraulically drive at least three of the plurality of actuated engine systems.

[0063] Hydraulically driving an actuated engine system may mean hydraulically driving at least one actuator within the actuated engine system.

[0064] The actuated engine systems may include one or more of the following: a thermal management system, a turbine casing cooling system, a generator thermal management system, an engine thermal management system, a bleed air system, a variable stator vane - VSV - system, a nacelle environmental control system, an anti - ice system, and a bearing chamber ventilation system.

[0065] The fuel supply system may be arranged to supply fuel to hydraulically drive at least four, five, six, or seven of the plurality of actuated engine systems.

[0066] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0067] The fuel supply system can be arranged to supply fuel to drive the turbine housing cooling system in a fuel hydraulic manner.

[0068] The engine can include an engine thermal management system. The fuel supply system can be arranged to supply fuel to drive the engine thermal management system in a fuel hydraulic manner.

[0069] The engine can include a generator thermal management system. The fuel supply system can be arranged to supply fuel to drive the generator thermal management system in a fuel hydraulic manner.

[0070] The engine can include a ventilation valve system. The fuel supply system can be arranged to supply fuel to drive the ventilation valve system in a fuel hydraulic manner.

[0071] The engine can include an air bleed system. The fuel supply system can be arranged to supply fuel to drive the air bleed system in a fuel hydraulic manner.

[0072] The spool can directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0073] The turbofan engine can include a gearbox that receives input from the spool and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the spool, such that the engine is a geared turbofan engine.

[0074] The engine can include a plurality of actuators. The fuel supply system can be arranged to supply fuel to drive at least ten, eleven, or twelve of the plurality of actuators in a fuel hydraulic manner.

[0075] At least one of the actuated engine systems can include at least two of the plurality of actuators. The fuel supply system can be arranged to supply fuel to drive at least one of the at least two actuators within the at least one actuated engine system in a fuel hydraulic manner.

[0076] At least one of the actuated engine systems can include at least two of the plurality of actuators. The fuel supply system can be arranged to supply fuel to drive each of the at least two actuators within the at least one actuated engine system in a fuel hydraulic manner.

[0077] The engine of the third aspect can be arranged to perform the method of the second aspect and can have any one of the features described with respect to the first or second aspect.

[0078] According to a fourth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0079] an engine core including a turbine, a combustor, a compressor and a spool connecting the turbine to the compressor;

[0080] a fan located upstream of the engine core and arranged to be driven by the spool (26), the fan including a plurality of fan blades;

[0081] a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outboard of the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0082] a plurality of actuated engine systems including a thermal management system and a turbine casing cooling system;

[0083] and a fuel supply system;

[0084] and wherein the method comprises:

[0085] supplying fuel using the fuel supply system for combustion in the combustor; and

[0086] supplying fuel using the fuel supply system to fuel hydraulically drive at least three of the plurality of actuated engine systems.

[0087] The method may comprise supplying fuel using the fuel supply system to fuel hydraulically drive at least four or five of the plurality of actuated engine systems.

[0088] The engine may include an engine thermal management system and a generator thermal management system. The method may comprise supplying fuel to fuel hydraulically drive one or both of the engine thermal management system and the generator thermal management system.

[0089] The methods of the second and fourth aspects may be complementary and may be carried out together in various embodiments. The method of the fourth aspect may be carried out using the engine of the first or third aspect.

[0090] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0091] an engine core including a turbine, a combustor, a compressor and a spool connecting the turbine to the compressor;

[0092] a fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0093] A nacelle that surrounds a fan and an engine core and defines a bypass duct located radially outward of the engine core, where the bypass ratio is at least 4, and 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 under cruise conditions;

[0094] A plurality of actuators, where at least one of the plurality of actuators is configured to actuate a bleed air valve; and

[0095] A fuel supply system, where the fuel supply system is arranged to supply fuel for combustion in a combustor and to supply fuel to hydraulically drive at least one actuator configured to actuate a bleed air valve.

[0096] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0097] Thus, the flow of bleed air can be controlled hydraulically by fuel. Bleed air typically comprises air "bled" from one or more compressors of the engine and / or from an auxiliary power unit (APU). An aircraft can include a bleed air system that includes a plurality of ducts, valves, and regulators arranged to direct the medium to high pressure air "bled" from the compressor and / or APU to various locations within the aircraft (within the engine and elsewhere). Bleed air can be used for a variety of different functions, including cabin pressurization, air conditioning, engine starting, bearing chamber ventilation, and de-icing. At least a portion of the aircraft bleed air system can form part of the engine or be contained within the engine, and this portion can be referred to as the engine bleed air system. The aircraft bleed air system can include a plurality of bleed air systems, some of which are entirely located within the engine.

[0098] At least one bleed air valve can be an ambient bleed air valve. Thus, bleed air can be used to pressurize the aircraft cabin and thereby supply air to the environmental control system. At least one bleed air valve can be a regulating component valve. As an example of an ambient bleed air valve, the component valve can regulate the bleed air flow into an air cycle air conditioning system.

[0099] Additionally or alternatively, bleed air can be used to keep critical parts of the aircraft (such as the leading edge of the wing, nacelle parts around the fan, and / or stator parts of the engine) from icing. The bleed air valve can be referred to as an anti-ice valve.

[0100] At least one bleed air valve can be a nacelle anti-ice valve.

[0101] At least one bleed air valve can be an engine part stator anti-ice valve.

[0102] The engine may include an isolation valve configured to enable the bleeding of air from an external engine input, i.e., the isolation valve may allow cross - supply of air from another engine of the same aircraft (it should be understood that for commercial aircraft, there are typically at least two engines - one per wing, and many aircraft have multiple engines per wing, e.g., a total of four or six engines. Thus, the bleed air can be shared between engines). At least one of the plurality of actuators may be configured to actuate the isolation valve. The fuel supply system may be arranged to supply fuel to hydraulically drive, via fuel, at least one actuator configured to actuate the isolation valve.

[0103] The engine may include an engine starter valve. At least one of the plurality of actuators may be configured to actuate the engine starter valve. The fuel supply system may be arranged to supply fuel to hydraulically drive, via fuel, at least one actuator configured to actuate the engine starter valve. Thus, bleed air drawn from the aircraft's auxiliary power unit (APU) or an operating engine can be used to power an air - turbine starter motor to start the aircraft's engine / another engine. Using an air - turbine starter may allow a smaller and lighter unit to provide the required torque compared to an electric or hydraulically driven starter. The fuel - hydraulic control of the engine starter valve may further reduce the size and / or weight of the system.

[0104] The engine may include an advanced valve for controlling the airflow from the compressor. At least one of the plurality of actuators may be configured to actuate the advanced valve. The fuel supply system may be arranged to supply fuel to hydraulically drive, via fuel, at least one actuator configured to actuate the advanced valve. The advanced valve may be configured to control the flow of high - pressure bleed air from the compressor to the aircraft cabin.

[0105] The engine may include an auxiliary power unit - APU - and an auxiliary power unit valve, or be associated therewith. In some embodiments, the APU may be separated from the engine and installed elsewhere on the aircraft. At least one of the plurality of actuators may be configured to actuate the auxiliary power unit valve. The fuel supply system may be arranged to supply fuel to hydraulically drive, via fuel, at least one actuator configured to actuate the auxiliary power unit valve.

[0106] The engine may include a manifold pressure valve. At least one of the plurality of actuators may be configured to actuate the manifold pressure valve. The fuel supply system may be arranged to supply fuel to hydraulically drive, via fuel, at least one actuator configured to actuate the manifold pressure valve. The manifold pressure valve may be configured to control the flow of high - pressure bleed air from the engine. The manifold pressure valve may regulate / control the pressure of the bleed air flow from the compressor to the aircraft cabin.

[0107] The fuel supply system may be arranged to supply fuel to hydraulically drive at least ten, at least eleven, or at least twelve of a plurality of actuators.

[0108] The engine may include a plurality of actuated engine systems. The plurality of actuated engine systems may include a thermal management system and a turbine casing cooling system. The fuel supply system may be arranged to supply fuel to hydraulically drive at least three of the plurality of actuated engine systems.

[0109] The actuated engine systems may include one or more of the following: a thermal management system, a turbine casing cooling system, a generator thermal management system, an engine thermal management system, a bleed air system, a variable stator vane - VSV - system, a nacelle environmental control system, an ice protection system, and a bearing chamber ventilation system.

[0110] The fuel supply system may be arranged to supply fuel to hydraulically drive at least four, five, six, or at least seven of the actuated engine systems.

[0111] The spool may directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0112] The turbofan engine may include a gearbox that receives input from the spool and outputs drive to the fan, thereby driving the fan at a lower rotational speed than the spool, such that the engine is a geared turbofan engine.

[0113] The engine of the fifth aspect may be arranged to perform the methods of the second and / or fourth aspects and may have any of the features described with respect to the first to fourth aspects.

[0114] According to a sixth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0115] An engine core that includes a turbine, a combustor, a compressor, and a spool that connects the turbine to the compressor;

[0116] A fan that is located upstream of the engine core and is arranged to be driven by the spool, the fan including a plurality of fan blades;

[0117] A nacelle that surrounds the fan and the engine core and defines a bypass duct that is radially outward of the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0118] A plurality of actuators, wherein at least one of the plurality of actuators is configured to actuate a bleed air valve; and

[0119] Fuel supply system;

[0120] And wherein the method comprises:

[0121] Supplying fuel using a fuel supply system for combustion in a burner; and

[0122] Supplying fuel using a fuel supply system to hydraulically drive, by the fuel, at least one actuator configured to actuate an air bleed valve.

[0123] The engine may include a plurality of air bleed valves. The method may include supplying fuel using a fuel supply system to hydraulically drive, by the fuel, a plurality of actuators each configured to actuate an air bleed valve.

[0124] In various embodiments, the methods of the second, fourth, and sixth aspects may be complementary and may be performed together, or in any combination or sub - combination. The method of the sixth aspect may be performed using the engine of the first, third, or fifth aspect.

[0125] According to a seventh aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0126] An engine core including a turbine, a burner, a compressor, and a spool connecting the turbine to the compressor;

[0127] A fan located upstream of the engine core and arranged to be driven by the spool (26), the fan including a plurality of fan blades;

[0128] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0129] An engine thermal management system;

[0130] A plurality of actuators including an actuator configured to actuate at least one valve within the engine thermal management system;

[0131] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive, by the fuel, an actuator configured to actuate at least one valve within the engine thermal management system;

[0132] Wherein the actuator configured to actuate at least one valve within the engine thermal management system is configured to actuate the valve so as to enable non - binary position adjustment between an open valve position and a closed valve position.

[0133] An actuator configured to actuate at least one valve within an engine thermal management system may be configured to actuate the valve to enable position adjustment between an open valve position, one or more discrete intermediate valve positions, and a closed valve position. Thus, non-binary position adjustment may include discrete position adjustment between at least three different valve positions. The actuator may be configured to actuate the valve between at least four set positions. For example, the actuator may be configured to actuate the valve between four set positions (open, closed, and two discrete / set intermediate positions).

[0134] An actuator configured to actuate at least one valve within an engine thermal management system may be configured to actuate the valve to enable continuous position adjustment between an open valve position and a closed valve position.

[0135] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0136] The engine thermal management system may include a plurality of valves.

[0137] An actuator configured to actuate at least one valve within an engine thermal management system may be configured to actuate at least two valves among the plurality of valves within the engine thermal management system.

[0138] The plurality of actuators may include a plurality of actuators each configured to actuate at least one valve among the plurality of valves within the engine thermal management system.

[0139] The engine may include a bleed air valve. The plurality of actuators may include an actuator configured to actuate the bleed air valve. The fuel supply system may be arranged to supply fuel to hydraulically drive the actuator configured to actuate the bleed air valve.

[0140] The engine may include a generator thermal management system. The plurality of actuators may include an actuator configured to actuate at least one valve within the generator thermal management system. The fuel supply system may be arranged to supply fuel to hydraulically drive the actuator configured to actuate at least one valve within the generator thermal management system.

[0141] The generator thermal management system may include a plurality of valves.

[0142] At least one actuator configured to actuate at least one valve within the generator thermal management system may be configured to actuate at least two valves among the plurality of valves within the generator thermal management system.

[0143] The plurality of actuators may include a plurality of actuators each configured to actuate at least one valve among the plurality of valves within the generator thermal management system.

[0144] An actuator configured to actuate at least one valve within a generator thermal management system may be configured to actuate at least one valve within the generator thermal management system so as to enable continuous position adjustment between an open valve position and a closed valve position.

[0145] The fuel supply system may be arranged to supply fuel to hydraulically drive at least ten, eleven, or at least twelve of the plurality of actuators by fuel.

[0146] The engine may include a plurality of actuated engine systems. The engine systems may include a thermal management system and a turbine housing cooling system. The engine systems may include a VSV actuation system, one or more de-icing systems, and / or a nacelle air conditioning system. The fuel supply system may be arranged to supply fuel to hydraulically drive at least three of the plurality of actuated engine systems by fuel.

[0147] The fuel supply system may be arranged to supply fuel to hydraulically drive at least four, five, six, or seven of the actuated engine systems by fuel.

[0148] The spool may directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0149] The turbofan engine may include a gearbox that receives input from the spool and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the spool, such that the engine is a geared turbofan engine.

[0150] The engine thermal management system may include an air-oil heat exchanger. An actuator configured to actuate at least one valve within the engine thermal management system may be configured to actuate the air-side valve of the air-oil heat exchanger. Accordingly, the actuator may control air flow. An actuator configured to actuate at least one valve within the engine thermal management system may be configured to actuate the oil-side valve of the air-oil heat exchanger. Accordingly, the actuator may control oil flow.

[0151] The engine thermal management system may include one or more, and optionally a plurality of, air-oil heat exchangers.

[0152] The engine thermal management system may include one or more, and optionally a plurality of, fuel-oil heat exchangers.

[0153] One or more, and optionally each, of the plurality of heat exchangers may include a valve controllable via an associated actuator. The fuel supply system may be arranged to supply fuel to hydraulically drive each of the associated actuators by fuel.

[0154] At least one valve within the engine heat management system can be configured such that oil can bypass one or more of the plurality of air-oil heat exchangers.

[0155] At least one valve within the engine heat management system can be configured such that air can bypass one or more of the plurality of air-oil heat exchangers.

[0156] At least one valve within the engine heat management system can be configured such that fuel can bypass one or more of the plurality of fuel-oil heat exchangers.

[0157] At least one valve within the engine heat management system can be configured such that oil can bypass one or more of the plurality of fuel-oil heat exchangers.

[0158] The fluid portion that can bypass the respective heat exchanger can be adjustable.

[0159] The engine of the seventh aspect can be arranged to perform the methods of the second, fourth, and / or sixth aspects and can have any of the features described with respect to the first to sixth aspects.

[0160] According to an eighth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0161] An engine core that includes a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor;

[0162] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0163] A nacelle that surrounds the fan and the engine core and defines a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0164] An engine heat management system;

[0165] A plurality of actuators, including an actuator configured to actuate a valve within the engine heat management system, wherein the actuator configured to actuate a valve within the engine heat management system is configured to actuate the valve so as to enable non-binary position adjustment between an open valve position and a closed valve position; and

[0166] A fuel supply system;

[0167] And wherein the method includes:

[0168] Supplying fuel using the fuel supply system for combustion in the combustor; and

[0169] Fuel is supplied using a fuel supply system to drive at least an engine thermal management system valve actuator among a plurality of actuators in a fuel hydraulic manner.

[0170] The position adjustment may be continuous or may be discrete between three or more different positions and optionally between four or more different positions.

[0171] In various embodiments, the methods of the second, fourth, sixth, and eighth aspects may be complementary and may be performed together or in any combination or sub - combination. The method of the eighth aspect may be performed using the engine of the first, third, fifth, or seventh aspect.

[0172] According to a ninth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0173] An engine core including a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor;

[0174] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0175] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0176] A generator thermal management system;

[0177] A plurality of actuators including an actuator configured to actuate at least one valve within the generator thermal management system; and

[0178] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the combustor and to supply fuel to drive the actuator, which is configured to actuate at least one valve within the generator thermal management system, in a fuel hydraulic manner;

[0179] wherein the actuator configured to actuate at least one valve within the generator thermal management system is configured to actuate the valve so as to enable non - binary position adjustment between an open valve position and a closed valve position.

[0180] An actuator configured to actuate at least one valve within a generator thermal management system may be configured to actuate the valve so as to enable position adjustment between an open valve position, one or more discrete intermediate valve positions, and a closed valve position. Thus, non-binary position adjustment may include discrete position adjustment between at least three different valve positions. The actuator may be configured to actuate the valve between at least four set positions. For example, the actuator may be configured to actuate the valve between four set positions (open, closed, and two intermediate positions).

[0181] An actuator configured to actuate at least one valve within a generator thermal management system may be configured to actuate the valve so as to enable continuous position adjustment between an open valve position and a closed valve position.

[0182] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0183] The generator thermal management system may be arranged to control the temperature of the generator by means of cooling oil for cooling and lubricating the generator, thereby transferring heat from the generator.

[0184] The generator thermal management system may include a plurality of heat exchangers; for example, one or more air-oil heat exchangers and one or more fuel-oil heat exchangers. The actuator may be arranged to control the fluid flow of the fluid through one or more of the heat exchangers, such as oil flow, air flow, or fuel flow.

[0185] The engine may include a plurality of generators. The generator thermal management system may be configured to manage the temperatures of all of the plurality of generators. The engine may include a plurality of generator thermal management systems, each system associated with a respective one of the plurality of generators.

[0186] The generator thermal management system may include a plurality of valves. At least one actuator configured to actuate at least one valve within the generator thermal management system may be configured to actuate at least two valves among the plurality of valves within the generator thermal management system.

[0187] The generator thermal management system may include a plurality of valves. The plurality of actuators may include a plurality of actuators configured to actuate at least one valve among the plurality of valves within the generator thermal management system.

[0188] The engine may include an ambient relief valve. The plurality of actuators may include an actuator configured to actuate the ambient relief valve. The fuel supply system may be arranged to supply fuel to hydraulically drive the actuator configured to actuate the ambient relief valve.

[0189] The engine may include an engine thermal management system. The plurality of actuators may include an actuator configured to actuate at least one valve within the engine thermal management system. The fuel supply system may be arranged to supply fuel to hydraulically drive, by fuel hydraulics, an actuator configured to actuate at least one valve within the engine thermal management system.

[0190] The engine thermal management system may include a plurality of valves.

[0191] At least one actuator configured to actuate at least one valve within the engine thermal management system may be configured to actuate at least two valves among the plurality of valves within the engine thermal management system.

[0192] The plurality of actuators may include a plurality of actuators each configured to actuate at least one valve among the plurality of valves within the engine thermal management system.

[0193] An actuator configured to actuate at least one valve within the engine thermal management system may be configured to actuate at least one valve within the engine thermal management system so as to enable continuous position adjustment between an open valve position and a closed valve position.

[0194] The fuel supply system may be arranged to supply fuel to hydraulically drive, by fuel hydraulics, at least ten, eleven, or at least twelve actuators among the plurality of actuators.

[0195] The engine may include a plurality of actuated engine systems. The engine actuation system may include a hydro-mechanical unit system and a turbine casing cooling system. The fuel supply system may be arranged to supply fuel to hydraulically drive, by fuel hydraulics, at least three actuated engine systems among the plurality of actuated engine systems.

[0196] The fuel supply system may be arranged to supply fuel to hydraulically drive, by fuel hydraulics, at least four, five, six, or seven actuated engine systems among the actuated engine systems.

[0197] The spool may directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0198] The turbofan engine may include a gearbox that receives an input from the spool and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the spool, such that the engine is a geared turbofan engine.

[0199] The engine of the ninth aspect may be arranged to perform the methods of the second, fourth, sixth, and / or eighth aspects and may have any one of the features described with respect to the first to eighth aspects.

[0200] According to a tenth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0201] An engine core, which includes a turbine, a combustor, a compressor, and a shaft connecting the turbine to the compressor;

[0202] A fan, which is located upstream of the engine core and is arranged to be driven by the shaft, and the fan includes a plurality of fan blades;

[0203] A nacelle, which surrounds the fan and the engine core and defines a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, and the bypass ratio is defined as the ratio of the mass flow rate of the flow passing through the bypass duct to the mass flow rate of the flow passing through the core under cruise conditions;

[0204] A generator thermal management system;

[0205] A plurality of actuators, including an actuator configured to actuate a valve within the generator thermal management system, wherein the actuator configured to actuate the valve within the generator thermal management system is configured to actuate the valve so as to be able to perform a non-binary position adjustment between an open valve position and a closed valve position; and

[0206] A fuel supply system;

[0207] And wherein the method includes:

[0208] Using the fuel supply system to supply fuel for combustion in the combustor; and

[0209] Using the fuel supply system to supply fuel to hydraulically drive at least the valve actuator of the generator thermal management system among the plurality of actuators.

[0210] The position adjustment can be continuous, or can be discrete between three or more different positions, and optionally between four or more different positions.

[0211] In various embodiments, the methods of the second, fourth, sixth, eighth, and tenth aspects can be complementary and can be performed together, or in any combination or sub-combination. The method of the tenth aspect can be performed using the engine of the first, third, fifth, seventh, or ninth aspect.

[0212] It should be understood that the features described with respect to one aspect can be used in combination with any other aspect, with necessary modifications.

[0213] As described elsewhere herein, the present disclosure can be applied to any relevant configuration of a gas turbine engine. Such gas turbine engines can be, for example, turbofan gas turbine engines, open rotor gas turbine engines (wherein the propeller is not surrounded by a nacelle), turboprop engines, or turbojet engines. Any such engine can be provided with or without an afterburner. Such gas turbine engines can be configured, for example, for land-based or marine power generation applications.

[0214] A gas turbine engine according to any aspect of the present disclosure may include an engine core including a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor. Such a gas turbine engine 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).

[0215] 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 configured 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.

[0216] 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.

[0217] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct drive turbofan engine in which the fan is 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, the compressor, and the 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.

[0218] 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 the spool and outputs a 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).

[0219] A gas turbine engine as described herein and / or claimed may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbines and compressors, such as one shaft, two shafts or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may also include a second turbine, a second compressor and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

[0220] In such an arrangement, the second compressor may be axially located downstream of the first compressor. The second compressor may be arranged to receive the flow (e.g. directly receive, e.g. via a substantially annular duct) from the first compressor.

[0221] The gearbox may be arranged to be driven by a core shaft (such as the first core shaft 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 core shaft configured (e.g. in use) to rotate at the lowest rotational speed (e.g., in the above example, only by the first core shaft and not the second core shaft). 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.

[0222] The gearbox may be a reduction gearbox (since the output to the fan has a lower rotational rate than the input from the core shaft). 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.

[0223] 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), such as greater than 2.5, such as in the range of 3 to 4.2, or 3.2 to 3.8, such as, 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 tooth 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 tooth ratio can be outside of these ranges.

[0224] 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.

[0225] The compressor or each compressor (e.g., the first compressor and the second compressor as described above) can include any number of stages, such as 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 (excluding 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 (excluding 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.

[0226] 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., regardless of 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.

[0227] Each fan blade may be defined as having a radial span that extends from a root (or hub) at a radially inner gas washing location or 0% span location to a tip at a 100% span location. 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 an 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 (or axially foremost) part 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.

[0228] 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 an 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.

[0229] 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.

[0230] When using a gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. 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).

[0231] 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.

[0232] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before entering the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example only, 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.

[0233] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, for a given thrust condition, the specific thrust can depend on the specific 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.

[0234] 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, the maximum thrust of a geared gas turbine engine having a fan diameter within the range of 200 cm to 210 cm may be within the range of 140 kN to 160 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine having a fan diameter within the range of 210 cm to 230 cm may be within the range of 150 kN to 200 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine having a fan diameter within the range of 340 cm to 360 cm may be within the range of 370 kN to 500 kN. By way of non-limiting example only, the maximum thrust of a direct drive gas turbine engine having a fan diameter within the range of 300 cm to 340 cm may be 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.

[0235] 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 itself may 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. 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 in the range of 1540K to 1600K under cruise conditions. 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 in the range of 1590K to 1650K under cruise conditions. 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 in the range of 1600K to 1660K under cruise conditions. 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 in the range of 1590K to 1650K under cruise conditions. 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 in the range of 1570K to 1630K under cruise conditions.

[0236] The TET under cruise conditions 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, from 1530K to 1600K. 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. 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 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. The maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0237] 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 a 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 resists impact (e.g., from birds, ice, or other materials) better than the rest of the blade. Such a leading edge 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.

[0238] As described and / or claimed herein, the fan 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). 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 to secure the fan blade to the hub / disc. 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 part of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0239] 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.

[0240] The fan of a gas turbine as described and / or claimed herein 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 metallic body (such as an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0241] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, landing, approach, and landing (or one or more parts thereof) have their conventional meanings and will be readily understood by a person skilled in the art. Thus, for a given gas turbine engine for an aircraft, a person skilled in the art will immediately recognize that each term refers to all or one or more parts of the operating phase of the engine within a given mission of the aircraft to which the gas turbine engine is designed to be attached.

[0242] 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 a further non-limiting example, the engine may produce between 5% and 8% of the available thrust. In a further 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 a further non-limiting example, the engine may produce between 6% and 12% of the available thrust. In a further 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 a further non-limiting example, the engine may produce between 95% and 100% of the available thrust. In a further non-limiting example, the engine may produce 100% of the available thrust.

[0243] 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 a further non-limiting example, the engine may produce between 80% and 95% of the available thrust. In a further non-limiting example, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to an operating phase during the aircraft flight cycle between takeoff and reaching cruise conditions, such that reaching cruise conditions defines the start of the cruise phase or a portion thereof of the aircraft flight. Additionally or alternatively, climb may refer to a nominal point or one or more nominal time periods 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.

[0244] As used herein, the cruise conditions that may define the cruise phase (or a portion thereof) of an aircraft flight have their conventional meaning and are 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 of the engine during the intermediate cruise of a given mission of the aircraft to which the gas turbine engine is designed to be attached (which may be referred to in the industry as the "economic mission"). In this regard, the intermediate cruise may be considered as the point in the aircraft flight cycle at which 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, the cruise conditions may define the operating point, phase or a portion thereof of the flight, which, taking into account the number of engines provided to the aircraft to which the gas turbine engine is designed to be attached, provides the thrust that will ensure the steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number) or at least substantially steady-state operation (i.e., maintaining at least substantially a constant altitude and / or at least substantially a constant Mach number) of the aircraft. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, then under cruise conditions, the engine may provide half of the total thrust required for the steady-state operation or at least substantially steady-state operation of the aircraft during intermediate cruise.

[0245] In other words, for a given gas turbine engine of an aircraft, the cruise conditions may be defined as the operating point of the engine that provides a specified thrust under intermediate cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the intermediate 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 intermediate cruise Mach number. For any given gas turbine engine of an aircraft, the intermediate cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine under cruise conditions can be well-defined.

[0246] By way of example only, the forward speed under cruise conditions may be any point in the range from 0.7 Mach to 0.9 Mach, such as from 0.75 to 0.85, such as from 0.76 to 0.84, such as from 0.77 to 0.83, such as from 0.78 to 0.82, such as from 0.79 to 0.81, such as approximately 0.8 Mach, approximately 0.85 Mach or in the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, such as below 0.7 Mach or above 0.9 Mach.

[0247] 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: 10000m to 15000m, for example within the range of 10000m to 12000m, for example within the range of 10400m to 11600m (about 38000 feet), for example within the range of 10500m to 11500m, for example within the range of 10600m to 11400m, for example within the range of 10700m (about 35000 feet) to 11300m, for example within the range of 10800m to 11200m, for example within the range of 10900m to 11100m, for example approximately 11000m. The cruise conditions may correspond to the standard atmospheric conditions at any given altitude within these ranges.

[0248] 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 35000ft (10668m). Under such cruise conditions, the engine can provide a known required net thrust level. The known required net thrust level of course depends on the engine and its intended application and can be, for example, a value within the range of 20kN to 40kN.

[0249] By way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Under such cruise conditions, the engine can provide a known required net thrust level. The known required net thrust level of course depends on the engine and its intended application and can be, for example, a value within the range of 35kN to 65kN.

[0250] In use, the gas turbine engine described and / or claimed herein can operate under cruise conditions defined elsewhere herein. Such cruise conditions can be determined by the cruise conditions of an aircraft (e.g., intermediate cruise conditions) on which at least one (e.g., 2 or 4) gas turbine engines can be installed to provide propulsion thrust.

[0251] Furthermore, those skilled in the art will immediately recognize that either or both of landing and approach refer to the operational phases between the cruise and landing of an aircraft during its flight cycle, in particular the approach forms part of the landing and take-off (LTO) phase. During either or both of landing and approach, the engine can produce between 0% and 50% of the available thrust. In further non-limiting examples, the engine can produce between 25% and 40% of the available thrust. In further non-limiting examples, the engine can produce between 30% and 35% of the available thrust. Additionally or alternatively, landing can refer to the nominal point in 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.

[0252] 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 an aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise condition according to this aspect may correspond to an operating point, phase, or a part thereof of the flight of the aircraft, as defined elsewhere herein.

[0253] According to one aspect, there is provided a method of operating a gas turbine engine as described herein and / or claimed. The operation may be carried out under any cruise condition (e.g., in terms of thrust, atmospheric conditions, and Mach number) that may be defined elsewhere herein.

[0254] 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 may include (or may be) an operation under any suitable condition (e.g., at an intermediate cruise of the aircraft), as defined elsewhere herein.

[0255] The skilled person will understand that features or parameters described in relation to any one of the above aspects may be applied to any other aspect unless mutually exclusive. Additionally, any feature or parameter included or described herein may be applied to any aspect and / or combined with any other feature or parameter included or described herein unless mutually exclusive.

[0256] Unless mutually exclusive, any parameter or value included or described herein may be applied to any one or more additional parameters and / or values included or described herein and / or combined therewith. For example, a first parameter or value (e.g., parameter A) included or described herein may be applied to any one or more additional parameters and / or values (e.g., parameter B; parameter C; and one or more of parameter D, etc.) included or described herein and / or combined therewith to express a product of their relationships. For example, the skilled person in the art will understand that in the case where parameter A and parameter B are disclosed separately, the product of their relationship may be expressed, as required, for example, as A / B, B / A, B*A, or any such further application, combination, or function of parameter A with respect to parameter B. BRIEF DESCRIPTION OF THE DRAWINGS

[0257] Embodiments will now be described, by way of example only, with reference to the drawings, in which:

[0258] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0259] Figure 2 is a close-up cross-sectional side view of an upstream portion of a geared gas turbine engine;

[0260] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine.

[0261] Figure 4 is a close-up cross-sectional side view of the upstream portion of a directly-driven gas turbine engine;

[0262] Figure 5 is a diagram of an aircraft with a propulsion system including two gas turbine engines;

[0263] Figure 6 is a schematic diagram showing a part of a fuel supply system including one fuel hydraulic actuator;

[0264] Figure 7 is a schematic diagram showing a part of a fuel supply system including two fuel hydraulic actuators;

[0265] Figure 8 is a schematic diagram showing a part of a fuel supply system including one fuel hydraulic actuator system;

[0266] Figure 9 is a schematic diagram showing a part of a fuel supply system including two fuel hydraulic actuator systems;

[0267] Figure 10 is a schematic diagram showing an example of a fuel hydraulic actuator system;

[0268] Figure 11 is a schematic diagram showing another example of a fuel hydraulic actuator system;

[0269] Figure 12 is a schematic diagram showing another example of a fuel hydraulic actuator system;

[0270] Figure 13 is a schematic diagram showing another example of an actuator driven in a fuel hydraulic manner;

[0271] Figure 14 is a schematic diagram showing another example of an actuator driven in a fuel hydraulic manner;

[0272] Figure 15 is a schematic diagram showing another example of an actuator driven in a fuel hydraulic manner;

[0273] Figure 16 is a schematic diagram showing another example of an actuator driven in a fuel hydraulic manner;

[0274] Figure 17 is a schematic diagram showing another example of an actuator driven in a fuel hydraulic manner;

[0275] Figure 18It is a schematic diagram showing another example of an actuator driven by fuel hydraulics;

[0276] Figure 19 It is a schematic diagram showing another example of an actuator driven by fuel hydraulics;

[0277] Figure 20 It is a schematic diagram showing another example of an actuator driven by fuel hydraulics;

[0278] Figure 21 It is a schematic diagram showing another example of an actuator driven by fuel hydraulics;

[0279] Figure 22 It is a schematic diagram showing an engine heat management system that includes an oil side valve operable using a fuel hydraulic actuator;

[0280] Figure 23 It is a schematic diagram showing an engine heat management system that includes an air side valve operable using a fuel hydraulic actuator;

[0281] Figure 24 It is a schematic diagram showing an engine heat management system including multiple heat exchangers;

[0282] Figure 25 It is a schematic diagram showing an alternative engine heat management system including multiple heat exchangers;

[0283] Figure 26 It is a schematic diagram showing another alternative engine heat management system including an air - oil heat exchanger and a fuel - oil heat exchanger;

[0284] Figure 27 It is a schematic diagram showing an example of a part of a fuel supply system that has a valve operable using a fuel hydraulic actuator;

[0285] Figure 28 It is a schematic diagram showing a generator heat management system that includes an oil side valve operable using a fuel hydraulic actuator;

[0286] Figure 29 It is a schematic diagram showing a generator heat management system that includes an air side valve operable using a fuel hydraulic actuator;

[0287] Figure 30 It is a schematic diagram showing an engine oil - generator oil heat exchanger that includes a valve operable via a fuel hydraulic actuator;

[0288] Figure 31 It is a flowchart showing an example method of operating a gas turbine engine;

[0289] Figure 32is a flow chart showing another example method of operating a gas turbine engine;

[0290] Figure 33 is a flow chart showing another example method of operating a gas turbine engine;

[0291] Figure 34 is a flow chart showing another example method of operating a gas turbine engine; and

[0292] Figure 35 is a flow chart showing another example method of operating a gas turbine engine. DETAILED DESCRIPTION

[0293] Figure 1 shows a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air intake 12 and a propulsive fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which 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 a reduction gearbox 30.

[0294] 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 are then expanded 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 reduction gearbox 30 is a reduction gearbox.

[0295] 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.

[0296] 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.

[0297] 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 shown. Four planet gears 32 are shown, 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.

[0298] In Figure 2 and Figure 3 the epicyclic gearbox 30 shown 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.

[0299] 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 to connect the gearbox 30 to the engine 10. By way of another 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 (e.g., 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 shown. 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

[0300] Accordingly, the present disclosure extends to gas turbine engines having any arrangement in the type of gearbox (e.g., stellar or planetary gears), the support structure, the input and output shaft arrangements, and the bearing positions.

[0301] Optionally, the gearbox may drive additional and / or alternative components (e.g., a medium-pressure compressor and / or a booster compressor).

[0302] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of interconnected shafts. By way of a further example, Figure 1 the gas turbine engine shown in

[0303] 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 outside the core engine nozzle 20. However, this is not restrictive, and any aspect of the present disclosure may also apply 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 that may be referred to as a mixed-flow nozzle. One or both nozzles (whether mixed or split) may have a fixed or variable area. Figure 4 A cross-sectional view of one such engine is shown in

[0304] 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 inlet 12, a propulsive fan 23, a medium pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a medium pressure turbine 19a, a low pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines the air inlet 12 and the exhaust nozzle 20.

[0305] In use, air entering the air inlet 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 medium pressure compressor 14, and the bypass airflow B passes through a bypass duct 22 to provide propulsive thrust. The medium pressure compressor 14 compresses the airflow A before delivering the air to the high pressure compressor 15 where further compression occurs.

[0306] 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 medium pressure turbine 19a, and the low pressure turbine 19 before being discharged through the nozzle 20, thereby driving the high pressure turbine, the medium pressure turbine, and the low pressure turbine to provide additional propulsive thrust. The high pressure turbine 17, the medium pressure turbine 19a, and the low pressure turbine 19 each drive the high pressure compressor 15, the medium pressure compressor 14, and the fan 23 respectively through suitable interconnecting shafts.

[0307] 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 turbines to the compressors and / or the fan.

[0308] Although the example described relates to a turbofan engine, 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.

[0309] The geometry of the gas turbine engine 10 and its components are 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 mutually perpendicular.

[0310] The fuel F supplied to the combustion apparatus 16 may include fossil-based hydrocarbon fuels such as kerosene. Thus, the fuel F may include molecules from one or more of the chemical families of normal alkanes, isoalkanes, cycloalkanes, and aromatic hydrocarbons. Additionally or alternatively, when blended with, mixed with, or replaced by an alternative fuel, the fuel F may include renewable hydrocarbons produced from biological or non-biological resources, also referred to as sustainable aviation fuel (SAF). In each provided embodiment, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganic substances, and metals.

[0311] SAF is understood by those skilled in the art 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 drawn from the atmosphere and / or captured from industrial processes; or various 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; alternative fuels from non-biological sources; jatropha; halophytes, and algae, rather than being synthesized from fossil-based hydrocarbons. SAF is understood not to include fossil fuels.

[0312] The functional performance of a given fuel composition or blend of the fuel F for a given task may be defined at least in part by the capabilities of the Brayton cycle of the fuel-servicing gas turbine engine 10. Parameters defining the functional performance may include, for example, specific energy; energy density; thermal stability; and emissions including gaseous and / or particulate matter. In this regard, particulate matter emissions may include soot particles produced by the combustion of the fuel F, also referred to as non-volatile particulate matter (nvPM). Any reference herein to soot or smoke may equally apply to other types of particulate matter emissions known in the art. Gaseous emissions may include any 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 gaseous emissions may equally apply to other types of gaseous emissions known in the art.

[0313] 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, thus potentially providing 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 can 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 spray breakup and atomization of the fuel.

[0314] 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 a 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 9 to 15 carbons and an average of 12 carbons.

[0315] 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), where additional compositions are 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.

[0316] Sustainable aviation fuel can include one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics and can be produced, for example, from one or more of: syngas; lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuel can include one or both of lower aromatic and sulfur content (relative to fossil-based hydrocarbon fuels). Additionally or alternatively, sustainable aviation fuel can include one or both of higher isoalkane and cycloalkane content (relative to fossil-based hydrocarbon fuels). In some examples, sustainable aviation fuel can include one or both of a density between 90% and 98% of kerosene density and a calorific value between 101% and 105% of kerosene calorific value.

[0317] In some examples, the aromatic and / or other non-alkane content of the sustainable aviation fuel or blend provided to the combustion equipment 16 may be relatively lower than that of kerosene. The sustainable aviation fuel may contain, for example, 30%, 20%, 15%, 10%, 8%, 5% or less than 5%; for example, 4%, 3%, 2%, 1% or less than 1%; for example, 0.75%, 0.5%, 0.25% or less than 0.25%; for example, 0.2%, 0.1% or less than 0.1%; for example, 0.01%, 0.001% or 0% of aromatic content. The aromatic content of the sustainable aviation fuel may be within an inclusive number or range defined by or within any two of the values in the previous sentence (i.e., the values may form an upper or lower limit), for example, 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; 0% to 0.75%, 0% to 0.5%, or 0.1% to 0.25%; 0.15% to 0.65%, 0.35% to 0.55% or 0.035% to 0.055%; determined according to one or more of preference, fuel feedstock or supplier, and compositional variations therein.

[0318] At least in part due to the molecular structure of the sustainable aviation fuel, the 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 equipment 16). Thus, relative to fossil-based hydrocarbon fuels such as kerosene, the sustainable aviation fuel may result in either or both a relative reduction in fuel consumption and a relative reduction in maintenance costs.

[0319] As Figure 5 shown, the aircraft 1 may include a plurality of fuel tanks 50, 53a, 53b; for example, a larger primary fuel tank 50 located in the aircraft fuselage and smaller fuel tanks 53a, 53b located in each wing. In other examples, the aircraft 1 may have only a single fuel tank 50, and / or the wing fuel tanks 53a, 53b may be larger than the central fuel tank 50, or the central fuel tank 50 may not be provided (instead, all fuel is stored in the wings of the aircraft) - it should be understood that many different tank layouts are conceivable, and the illustrated examples are provided for ease of description and are not intended to be limiting.

[0320] Figure 5An aircraft 1 having a propulsion system 2 including two gas turbine engines 10 is shown. Fuel is supplied to the gas turbine engines 10 from a fuel supply system on the aircraft 1. The fuel supply system illustrated in the figure includes a single fuel source. For the purposes of this application, the term "fuel source" means: 1) a single fuel tank; or 2) a plurality of fuel tanks that are fluidly interconnected. Each fuel source is arranged to provide a separate fuel source (i.e., the first fuel source may contain a first fuel that has one or more characteristics different from the second fuel contained in the second fuel source). Thus, the first fuel source and the second fuel source are not fluidly coupled to each other in order to separate different fuels (at least under normal operating conditions). Using multiple fuel sources allows the aircraft 1 to carry multiple different fuels and, during operation, optionally even switch between fuels during cruise or between different operating phases during flight.

[0321] In this example, the first (and only in these examples) fuel source includes a central fuel tank 50 located primarily in the fuselage of the aircraft 1 and a plurality of wing fuel tanks 53a, 53b, where at least one wing fuel tank is located in the left wing and at least one wing fuel tank is located in the right wing for balance. In the example shown, all of the tanks 50, 53a, 53b are fluidly interconnected, thereby forming a single fuel source. Each of the central fuel tank 50 and the wing fuel tanks 53a, 53b may include a plurality of fluidly interconnected fuel tanks.

[0322] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the central tank 50, thereby forming a separate second fuel source. For balance purposes, one or more fuel tanks in the left wing may be fluidly connected to one or more fuel tanks in the right wing. This may be done via the central fuel tank (if the fuel tank is not part of another fuel source) or bypassing the central fuel tank or both (for maximum flexibility and safety). In another example, the first fuel source includes the wing fuel tanks 53 and the central fuel tank 50, while the second fuel source includes another separate central fuel tank. Fluid interconnectivity may be provided between the wing fuel tanks and the central fuel tank of the first fuel source for balancing the aircraft 1. In an aircraft 1 having multiple fuel sources, two or more of the fuel sources may thus contain fuels that are different from each other such that the aircraft 1 can change fuels during flight. Thus, determining the fuel supplied to the burner 16 may be more complex than simply recording a single identifier of the fuel on the aircraft 1 or checking it once at startup.

[0323] In some examples, the distribution of fuel tanks 50, 53 available on the aircraft 1 may be restricted such that the first and second fuel sources are each substantially symmetric with respect to the aircraft centerline. In cases where asymmetric fuel tank distribution is allowed, suitable fuel transfer devices are typically provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source such that the position of the aircraft's center of mass can be maintained within acceptable lateral limits throughout the flight.

[0324] Aircraft typically refuel at multiple different airports, for example at the start and end of a long - haul flight. While there are standards that all aviation fuels must meet, as described above, different aviation fuels have different compositions, for example depending on their source (e.g., different petroleum sources, biofuels or other synthetic aviation fuels (commonly described as sustainable aviation fuel - SAF) and / or mixtures of petroleum - based fuels, and other fuels) and any additives included (e.g., such as antioxidants and metal deactivators, biocides, static reducers, icing inhibitors, corrosion inhibitors) as well as any impurities. In addition to differences between airports and fuel suppliers, even for a given airport or fuel supplier, the fuel composition of the available aviation fuel may vary between batches. Furthermore, the fuel tanks 50, 53 of the aircraft 1 are typically not emptied before being filled for a subsequent flight, resulting in a mixture of different fuels within the tanks - effectively a fuel with a different composition resulting from the mixture. One or more fuel ports 62 may be provided for refueling.

[0325] The fuel supply system 152 includes a main fuel flow path 110 from the fuel tank 50 to the burner 16 of the gas turbine engine 10, and various branch fuel paths that obtain fuel from or return fuel to the main fuel flow path 110.

[0326] Figure 6 A schematic view of a portion of the fuel supply system 152 of the gas turbine engine 10 is shown. The main fuel flow path 110 of the fuel supply system 152 serves as a fuel source for the fuel hydraulic system. The gas turbine engine 10 includes a burner 16 and an actuator 154. Figure 6 The dashed arrows in [figure] indicate the direction of fuel flow.

[0327] The fuel supply system 152 is arranged to supply fuel for combustion in the burner 16.

[0328] It should be understood that, in addition to other engine components, one or more valves, pumps, sensors, discharge devices for fuel hydraulic actuators, and heat exchangers may be present along the main fuel flow path 110. The fuel is supplied directly to the burner 16 via line 110a after passing through any preceding engine component furthest downstream along the main fuel flow path 110.

[0329] The fuel traveling through lines 159 and 160 is used to drive (i.e., actuate) actuator 154. Thus, actuator 154 is hydraulically driven by fuel. As used herein, the term "fuel hydraulic" refers to the hydraulic operation of an actuator that uses a hydraulic fluid as fuel. An actuator hydraulically driven by fuel may be referred to herein as a "fuel hydraulic actuator".

[0330] Fuel travels from fuel flow path 110 through line 159 to actuator 154. After being used to hydraulically drive actuator 154, the fuel returns via line 160 to the main fuel flow path 110 and may then be delivered to burner 16.

[0331] Fuel return line 160 may include a valve configured to regulate the fuel flow returning to main fuel path 110, and the valve may be operated using a fuel hydraulic actuator. Fuel return line 160 may direct the fuel to a location along main fuel flow path 110, either before or after one or more heat exchangers in main fuel flow path 110.

[0332] Once the fuel reaches burner 16, it burns to provide thrust, as explained with respect to Figure 1 as explained.

[0333] Figure 6 A single fuel hydraulic actuator 154 on fuel flow circuits 159, 160 is shown. In an alternative embodiment, multiple fuel hydraulic actuators 154 may be located on a single fuel flow circuit 159, 160, using fuel taken from a single point along main fuel flow path 110. The multiple actuators may be arranged in series with respect to the fuel flow (such that all of the fuel in discharge line 159 passes through all of the actuators in sequence) or in parallel with respect to the fuel flow (with discharge line 159 branching and having one or more actuators on each branch).

[0334] Figure 7 A schematic view of an alternative portion of fuel supply system 152 of gas turbine engine 10 is shown. Fuel supply system 152 includes fuel sources 50, 53 (which may be a single fuel tank or multiple fuel tanks) and fuel flow path 110 between fuel sources 50, 53 and burner 16. The fuel supply system 152 shown includes two actuators 254a, 254b, each actuator being located on a fuel discharge return circuit separate from main fuel flow path 110. Figure 7 The dashed arrows in indicate the direction of fuel flow.

[0335] The two actuators 254a, 254b are hydraulically actuated by fuel. That is, both actuators use fuel from the fuel sources 50, 53 as the hydraulic fluid for hydraulic actuation. The fuel travels from the main fuel flow path 110 to the actuators 254a, 254b via the pipelines 259a, 259b respectively. After the fuel is used to drive the actuators 254a, 254b, it returns to the main fuel flow path via the pipelines 260a, 260b.

[0336] Figure 7 Only two actuators 254a, 254b are shown, but in other embodiments, any number of actuators can be hydraulically actuated by fuel. In some embodiments, the gas turbine engine 10 can include multiple actuators hydraulically actuated by fuel and multiple non-fuel hydraulically actuated actuators.

[0337] Figure 7 A single fuel hydraulic actuator 254 is shown on each fuel flow loop. In an alternative embodiment, multiple fuel hydraulic actuators 254 can be located on one or more of the multiple fuel flow loops. The multiple actuators can be arranged in series with respect to the fuel flow (such that all the fuel in the discharge pipe 159 passes through all the actuators in sequence) or in parallel with respect to the fuel flow (with the discharge pipe 159 branching and having one or more actuators on each branch).

[0338] For some embodiments including multiple fuel hydraulic actuators, all the fuel used to hydraulically actuate the actuators can be taken from a single take-off point on the main fuel flow path (i.e., they can all be on a single fuel flow loop). In other embodiments, the fuel supplied to one or more of the multiple fuel hydraulic actuators can come from different take-off points on the main fuel flow path (i.e., they can be on different fuel flow loops), but the fuel from each take-off point will be at the same pressure.

[0339] In various embodiments, the gas turbine engine 10 includes at least ten, eleven, twelve, thirteen, fourteen, fifteen, sixteen, seventeen or eighteen fuel hydraulic actuators. In some embodiments, each actuator can have its own dedicated fuel discharge pipes 159, 259 on the main fuel flow path 110 such that it has its own fuel flow loop. In other embodiments, two or more of these actuators can be located on the same fuel flow loop outside the main fuel flow path 110.

[0340] In various embodiments, the gas turbine engine 10 includes one or more actuated engine systems. Each actuated engine system may include one or more individual actuators. An "actuated engine system" (also referred to as an "actuated system") refers to a set of actuators that are used in combination to control a particular system within the engine 10 or perform a particular function.

[0341] For example, the actuated engine systems of the engine 10 may include one or more of the following:

[0342] · An engine heat management system (HMS);

[0343] · A generator HMS;

[0344] · A variable stator vane (VSV) system;

[0345] · A nacelle environmental control system (e.g., a nacelle air conditioning system);

[0346] · An ice protection system;

[0347] · A turbine case cooling (TCC) system; and / or

[0348] · A bearing chamber ventilation system.

[0349] In addition, a given engine 10 may include multiple actuated systems of a given type, e.g., multiple ice protection systems each arranged to de-ice different components, multiple bearing chamber ventilation systems for different bearing chambers, and / or redundant nacelle environmental control systems.

[0350] Figure 8 A schematic diagram showing a portion of the fuel supply system 152 of the gas turbine engine 10 is shown. The fuel supply system 152 includes a main fuel flow path 110 from fuel sources 50, 53 to the burner 16, and an actuated system 356. Figure 8 The dashed arrows in [FIGURE REFERENCE] indicate the direction of fuel flow.

[0351] Similar to Figure 6 and Figure 7 , fuel is supplied to the burner 16 and fuel is supplied to the fuel hydraulic actuators 354b, 354c.

[0352] In Figure 8 's embodiment, the actuated system 356 includes three actuators 354a - 354c. In other embodiments, the actuated system 356 may include one actuator or any number of actuators.

[0353] In Figure 8In an embodiment, two actuators 354b, 354c are driven in a fuel hydraulic manner. Each fuel hydraulic actuator 354b, 354c of this embodiment has its own dedicated fuel flow circuit. In other embodiments, the fuel supplied to multiple actuators is taken from the same discharge point on the main fuel flow path 110.

[0354] The remaining actuator 354a is not driven in a fuel hydraulic manner, but can be driven by any other suitable means (e.g., hydraulic drive using oil, or electronic actuation, or pneumatic actuation. The dotted line represents any actuation means; it is a non-fuel liquid flow or an electrical signal).

[0355] In some embodiments, only one actuator within a given actuation system is driven in a fuel hydraulic manner. In other embodiments, any number of actuators within a given actuation system are driven in a fuel hydraulic manner. In some embodiments, all actuators within a given actuation system can be driven in a fuel hydraulic manner. Thus, any actuation system having at least one fuel hydraulic actuator can be referred to as a "fuel hydraulic" actuation system.

[0356] Figure 9 A schematic diagram showing a part of the fuel supply system 152 of the gas turbine engine 10 is presented. The fuel supply system 152 is generally as Figure 8 shown, but includes two actuated systems 456a, 456b. Figure 9 The dotted arrows in

[0357] still indicate the direction of fuel flow. The first actuated system 456a includes three actuators 454a - 454c, two of which are fuel hydraulic actuators. The second actuated system 456b includes two actuators 454d - 454e, both of which are fuel hydraulic actuators. The inclusion of this embodiment is for illustrative purposes only, and those skilled in the art will recognize that the turbine engine 10 can include any number of actuated systems, and each actuated system can include any number of actuators. Any number of these actuators can be driven in a fuel hydraulic manner.

[0358] The actuators described above can be any actuators used in the engine 10. Various specific actuators that can be driven in a fuel hydraulic manner are described below. The actuation systems described herein can be any actuation systems used in an aircraft. Various specific actuation systems that can be driven in a fuel hydraulic manner are described below. It should be understood that the specific embodiments listed are described only as examples, and the scope of protection is defined only by the claims.

[0359] An example of an actuated system is a variable stator vane (VSV) system. The VSV system controls the airflow through compressors 14, 15 (by moving one or more vanes; and typically by moving multi-stage compressor vanes) to provide optimal compressor performance. The VSV system changes the angle of the stator vanes of the compressor to manage the operability and efficiency of the compressor.

[0360] Typically, the variable stator vanes are arranged to be rotatable about an axis to align with the angle of the incoming airflow. This movement of the vanes allows a wider range of dimensionless flow through compressors 14, 15 without stalling; thanks to the VSV, compressors 14, 15 can operate effectively over a wider range of engine power settings without surging. The VSV is controlled by the engine control system in response to a series of parameters, which typically include environmental conditions and the thrust level required by the engine.

[0361] In an engine 10 having multiple compressors 14, 15, a VSV system may be provided for one, some, or all of the compressors.

[0362] Many VSV systems include two VSV actuators, which can be used to change the angle of the inlet guide vanes of compressors 14, 15. Each of the two VSV actuators can be configured to control the guide vanes on a given side of the engine 10. In various embodiments, one or more of the actuators of the VSV system are hydraulically driven by fuel.

[0363] Figure 10 A schematic diagram of a fuel supply system 152 of an aircraft is shown, which system includes a VSV system 556. The VSV system 565 includes two actuators 554a, 554b. The fuel supply system 152 is arranged to supply fuel to the burner 16 and hydraulically drive the actuators 554a, 554b by fuel.

[0364] Another example of an actuated system is a turbine casing cooling (TCC) system. A turbine engine 10 typically includes a casing surrounding turbines 17, 19, and the TCC system is used to selectively cool the casing. This cooling can be used to control the rotor-stator tip clearance by reducing the amount of thermal expansion. The TCC system includes one or more valves, which can be used to regulate the bleed airflow from the compressors 17, 19 of the engine through the casing to provide cooling.

[0365] In an engine 10 having multiple turbines 17, 19, a TCC system may be provided for one, some, or all of the turbines.

[0366] One or more valves of the TCC system can be operated via an actuator. In various embodiments, one or more of the actuators of the TCC system are hydraulically driven by fuel.

[0367] Figure 11 Figure 11 shows a schematic diagram of a fuel supply system 152 of an aircraft 1, which system includes a TCC system 656. The TCC system 656 includes two actuators 654, 654', each actuator 654, 654' being configured to control a corresponding valve 654a, 654a'. The fuel supply system 152 is arranged to supply fuel from fuel sources 50, 53 to a burner 16 and to hydraulically drive the actuators 654, 654' with the fuel.

[0368]

[0368] Another example of an actuated system used within a turbofan engine 10 is a hydromechanical unit (HMU). The HMU primarily distributes and regulates the fuel flow to the burner 16. It can also supply fuel to the actuators.

[0369]

[0369] The HMU includes a plurality of actuators configured to operate a plurality of valves. In various embodiments, one or more of these actuators are hydraulically driven with the fuel.

[0370]

[0370] The hydromechanical unit typically includes 3 or 4 actuators. In various embodiments, one, some, or all of these actuators can be hydraulically driven with the fuel.

[0371]

[0371] The hydromechanical unit can include an actuator configured to regulate the fuel flow to the burner 16 (e.g., it can be based on parameters including engine thrust demand and fuel calorific value), including controlling the fuel spillage around one or more engine fuel pumps (not all of the fuel leaving the pump has to be supplied to the burner 16; rather, some fuel can be recycled ("spilled"), and the recycled fuel typically forms part of the fuel entering the pump inlet. Thus, the spilled fuel can be used to perform engine functions as well as to allow the pump to operate at a set flow rate even when there are fluctuations in the burner fuel demand. Thus, as used herein, "spilled" fuel includes fuel used for any purpose other than being supplied to the burner 16, such as for fuel hydraulics, and not just fuel that is sent directly from the pump outlet to the pump inlet).

[0372] The hydraulic mechanical unit may include an actuator configured to cut off the fuel supply. The hydraulic mechanical unit may include an actuator configured to ensure a minimum fuel pressure or control the fuel pressure drop. The hydraulic mechanical unit may include an actuator configured to control the fuel supply through multiple pumps. The hydraulic mechanical unit may include an actuator configured to open a path to the drain manifold when the engine is shut down (e.g., a procedure for fuel return to the fuel tank, which may utilize a small auxiliary fuel tank installed on the engine 10 instead of the main fuel tanks 50, 53 of the aircraft). The hydraulic mechanical unit may include an actuator associated with VSV control. The hydraulic mechanical unit may include an actuator associated with thrust control adjustment. The hydraulic mechanical unit may include an actuator configured to split fuel distribution (e.g., distribute fuel to different manifolds to supply injectors to the burner 16, and / or for different fuel hydraulic controls). In some embodiments, TCC control and / or variable bleed valve control may be performed using one or more actuators located within the HMU, and thus there may be some overlap of the systems being actuated.

[0373] In some embodiments, the engine 10 may include a hydraulic mechanical metering unit that may be positioned upstream of a fuel metering valve that controls the fuel flow rate into the burner 16 along a main fuel flow path 110 through the engine. In some such embodiments, the fuel drain point or each fuel drain point may be located near or within the hydraulic mechanical metering unit.

[0374] Figure 12 A schematic diagram of a fuel supply system 152 of an aircraft 1 is shown, which system includes an HMU 756. The illustrated HMU 756 includes three actuators 754, 754’, 754”. The fuel supply system 152 is arranged to supply fuel from fuel sources 50, 53 to the burner 16 and to hydraulically drive the actuators 754 with fuel. Each actuator 754 is configured to control an associated valve 754a, 754a’, 754a”. Figure 12 Each actuator 754 is shown having its own dedicated drain point on the main fuel flow path 110 and having its own dedicated fuel circuit. In other embodiments, branch paths and / or distribution manifolds may be provided to split the fuel from a single drain point on the main fuel flow path 110 between multiple fuel hydraulic actuators. The hydraulic mechanical metering unit (if present) may be provided as part of the HMU.

[0375] Another example of an actuator used within the turbine engine 10 is a ventilation valve actuator (also referred to herein as a "vent valve actuator"). The vent valve actuator controls the opening and closing of a ventilation valve, which controls the opening and closing of a ventilation line for one or more engine bearing compartments.

[0376] Various embodiments may include one or more ventilation valves controlled by one or more ventilation valve actuators. In some embodiments, one or more of these ventilation valve actuators are fuel hydraulically actuated.

[0377] Figure 13 A schematic illustration of a fuel supply system 152 of the aircraft 1 is shown, which system includes a ventilation valve 854a. The ventilation valve 854a is controlled by an actuator 854. The fuel supply system 152 is arranged to supply fuel from fuel sources 50, 53 to fuel hydraulically actuate the actuator 854 and supply fuel to the burner 16. Actuation of the ventilation valve 854a can control the air flow between the engine bearing compartment and the atmosphere. A ventilation valve system can be defined that includes at least the ventilation valve 854a and its actuator 854, and optionally includes a plurality of valves and / or actuators.

[0378] Another example of an actuated system used within the turbine engine 10 is a bleed air system. Bleed air systems 954, 954a, 1054, 1054a, 1154, 1154a each including at least one valve and its associated actuator can redirect bleed air (compressed air taken from compressors 14, 15 upstream of the burner 16) for use in other systems. Bleed air is useful in various systems due to its relatively high temperature and pressure. The bleed air system can be referred to as an "engine bleed air system" (EBAS).

[0379] There are many bleed air systems that can be used within the turbine engine 10, some of which are described below. The bleed air systems include one or more valves controlled by one or more actuators. In various embodiments, one or more of the actuators within the bleed air system are fuel hydraulically actuated.

[0380] A bleed air system for controlling the pressure within the cabin of the aircraft 1 is referred to as an environmental bleed air system.

[0381] The bleed air system can remove or prevent ice formation on the nacelle 21 by redirecting high pressure and high temperature air from the compressors 14, 15 to the nacelle 21. Such systems include a nacelle anti-ice valve 954a controlled by an actuator 954. In various embodiments, the actuator is fuel hydraulically actuated.

[0382] Figure 14A schematic view of a fuel supply system 152 is shown, which includes fuel sources 50, 53 (combined within fuel flow path 110 for clarity) and a burner 16. The fuel supply system 152 is arranged to supply fuel to the burner 16 and to supply fuel to an actuator 954 so as to drive the actuator 954 in a fuel hydraulic manner. The actuator 954 is arranged to operate a valve 954a, which allows bleed air to pass through the nacelle 21 of the gas turbine engine 10.

[0383] The bleed air system may also remove or prevent ice formation on the engine stator by redirecting high-pressure and high-temperature air from compressors 14, 15 to the engine stator. Such systems may include an engine section stator (ESS) anti-ice valve controlled by an actuator. In various embodiments, the actuator is driven in a fuel hydraulic manner.

[0384] Figure 15 A schematic view of a fuel supply system 152 is shown, which includes fuel sources 50, 53 and a burner 16. The fuel supply system 152 is arranged to supply fuel to the burner 16 and to supply fuel to a bleed air actuator 1054 so as to drive the actuator 1054 in a fuel hydraulic manner. The actuator 1054 is configured to operate a valve 1054a, which allows bleed air to pass through and / or reach one or more stators of the turbine engine 10. The valve 1054a may be an ESS anti-ice valve.

[0385] The bleed air system may also be used to control the air conditioning within the cabin of the aircraft 1. In such systems, a conditioning assembly (either conditioning assembly or pressurized air conditioning kit (PACK) are acceptable terms) filters and cools the bleed air (via expansion and / or heat exchange with, for example, atmospheric air). The air conditioning assembly is used to cool the bleed air from the engine or APU and, where appropriate, dehumidify it, and then supply it to the aircraft cabin. Such systems include one or more valves and one or more actuators. Air from the engine 10 is typically supplied to the assembly via a one-way check valve 1154a. In various embodiments, one or more of these actuators are driven in a fuel hydraulic manner.

[0386] Figure 16 A schematic view of a fuel supply system 152 is shown, which is arranged to supply fuel to the burner 16 and to supply fuel to a conditioning assembly actuator 1154 so as to drive the actuator 1154 in a fuel hydraulic manner. The actuator 1154 is configured to operate a valve 1154a, which allows bleed air to enter the conditioning assembly. Thus, the valve 1154a may be referred to as an ambient bleed air valve; a conditioning assembly valve is an example of such a valve.

[0387] Before being used in other systems, bleed air is typically extracted from the compressors used in air conditioners through a high-level valve (also known as the "high-pressure valve") controlled by an actuator. In various embodiments, the actuator is driven in a fuel-hydraulic manner. The bleed air extracted from compressors 14, 15 via the high-level valve is then redirected via one or more additional valves for environmental control of the aircraft cabin.

[0388] Figure 17 A schematic diagram of a fuel supply system 152 is shown, which is arranged to supply fuel to a burner 16 and a high-level valve actuator 1254 so as to drive the actuator 1254 in a fuel-hydraulic manner. The actuator 1254 is configured to operate a high-level valve 1254a, which enables bleed air to be extracted from compressors 14, 15 for subsequent use.

[0389] In addition to or instead of compressors 14, 15, bleed air can be extracted from other components of the engine 10 and / or other components of the aircraft 1. For example, bleed air can be extracted from a bypass duct 22 or an APU. The extraction of this bleed air is carried out through an engine bleed air valve (also known as the "manifold pressure valve" or pressure regulating valve) controlled by an actuator. In various embodiments, the actuator is driven in a fuel-hydraulic manner.

[0390] Bleed air can typically be taken from three sources: intermediate pressure bleed (IP compressor 14), high pressure bleed (HP compressor 15), or bypass duct 22 (i.e., downstream of the fan 23).

[0391] Each bleed air source can have an associated valve. Additionally, there are typically valves for regulating the static pressure (downstream of the IP and HP sources) to meet the end-user requirements of the aircraft system. Any of these valves can be actuated in a fuel-hydraulic manner.

[0392] The manifold pressure valve is used to regulate the pressure of the bleed air to suit its intended use.

[0393] Figure 18 A schematic diagram of a fuel supply system 152 is shown, which is arranged to supply fuel to a burner 16 and an actuator 1354 so as to drive the actuator 1354 in a fuel-hydraulic manner. The actuator 1354 is configured to operate a valve 1354a, which enables bleed air to be extracted from engine components other than the compressor (e.g., from the bypass duct 22).

[0394] Bleed air can be extracted from an auxiliary power unit (APU) 1455; in some embodiments, the APU can be located within the engine 10, or in other embodiments at other locations within the aircraft. The extraction of bleed air from the APU is typically carried out through an APU bleed air valve controlled by an actuator. In various embodiments, the actuator is driven in a fuel-hydraulic manner.Figure 19 The dotted arrow therein indicates the bleed air flow from the APU 1454’ to the APU bleed air valve 1454a.

[0395] In some embodiments, the fuel for driving the APU bleed air valve may be taken from the supply line between the fuel source and the engine, while the fuel for other fuel hydraulic actuators may be taken from the supply line within the engine. Thus, in some embodiments, the APU bleed air valve 1454a and in fact the APU 1454’ may not be part of the engine 10.

[0396] Figure 19 A schematic diagram of a fuel supply system 152 is shown, which is arranged to supply fuel to the burner 16 and the APU bleed air actuator 1454 so as to drive the actuator 1454 in a fuel hydraulic manner. The actuator 1454 is configured to operate the valve 1454a, which enables bleed air to be drawn from the auxiliary power unit.

[0397] The bleed air from the APU 1455 can be used to assist various tasks, such as starting the engine 10. The engine start valve controls the bleed air flow from the APU 1455 to the turbines 17, 19. The engine start valve is controlled by an actuator. In various embodiments, the actuator is driven in a fuel hydraulic manner.

[0398] Figure 20 A schematic diagram of a fuel supply system 152 is shown, which is arranged to supply fuel to the burner 16 and the engine start valve actuator 1554 so as to drive the actuator 1554 in a fuel hydraulic manner. The actuator 1554 is configured to operate the engine start valve 1554a, which redirects the bleed air from the auxiliary power unit for starting the turbine engine 10.

[0399] For an aircraft having multiple engines 10 (e.g., on opposite wings of the aircraft 1), the bleed air taken from one turbine engine 10 can be supplied to the components of another turbine engine. The bleed air flow between the engines 10 is regulated by the isolation valve 1654a. The isolation valve is controlled by the actuator 1654. In various embodiments, the actuator is driven in a fuel hydraulic manner.

[0400] Figure 21 A schematic diagram of a fuel supply system 152 is shown, which is arranged to supply fuel to the burner 16 and the isolation valve actuator 1654 so as to drive the actuator 1654 in a fuel hydraulic manner. The actuator 1654 is configured to operate the isolation valve 1654a.

[0401] Another example of an actuation system used within the turbine engine 10 is the heat management system (HMS). The turbine engine 10 typically includes an engine HMS and a generator HMS. Both HMS systems control the temperature of the oil by enabling and controlling the use (or bypass) of an air - oil (and / or fuel - oil) heat exchanger.

[0402] The engine HMS system controls the temperature of the oil system that lubricates engine components (e.g., bearing housings, gearboxes, etc.). The engine HMS system includes one or more engine HMS valves controlled by one or more actuators. In various embodiments, one or more of these actuators are fuel - hydraulically driven.

[0403] For example, in some embodiments, the engine HMS includes a single valve and a single actuator that is fuel - hydraulically driven and arranged to control the valve.

[0404] The engine HMS system may include multiple heat exchangers arranged in a parallel configuration or a series configuration. A single valve controlled by a single fuel - hydraulic actuator enables the use (or bypass) of all heat exchangers. In particular, the flow rate of the fluid (air, oil, or fuel, if applicable) can be controlled by the valve, and the portion of the fluid that does not pass through the heat exchanger is directed along a bypass pipe. In some embodiments, each heat exchanger has an associated valve controlled by an associated actuator. In various embodiments, one or more of these actuators are fuel - hydraulically driven. Additionally, one or more recirculation pipes may be provided to allow the fluid to return from the heat exchanger outlet to the heat exchanger inlet, thus passing through the heat exchanger multiple times. The flow through the recirculation pipe can again be regulated using a valve controlled by an actuator. In some embodiments, a pump or other device for delivering fuel may be present within or associated with the recirculation pipe to drive this recirculation.

[0405] Figure 22 A schematic diagram showing a portion of the engine HMS 1700 is presented. The engine HMS includes an oil - air heat exchanger 1703 configured to transfer heat from the oil within the engine. The engine HMS 1700 includes an oil line 1701 and an air line 1702.

[0406] The engine HMS 1700 includes a valve 1704a positioned within the oil line 1701. The valve 1704a is configured to direct the oil through the heat exchanger 1703 or around the heat exchanger 1703. The valve 1704a is controlled by an actuator 1704 that is fuel - hydraulically driven.

[0407] Actuator 1704 is configured to actuate valve 1704a to enable non-binary position adjustment between an open valve position and a closed valve position (i.e., actuator 1704 can simulate rather than binary control valve 1704a, or it can enable the valve to be adjusted to one or more intermediate positions between open and closed). Thus, valve 1704a and associated fuel hydraulic actuator 1704 are configured to control the amount of oil passing through heat exchanger 1703 to regulate the cooling of the oil.

[0408] Figure 23 A schematic illustration of a portion of engine HMS1700 according to an alternative embodiment is shown. Engine HMS1700 includes an oil-air heat exchanger 1703 configured to transfer heat from oil within the engine. Engine HMS1700 includes an oil line 1701 and an air line 1702, as described in the Figure 22 described embodiment. However, the position of valve 1704a' controlled in a fuel hydraulic manner is different.

[0409] Figure 23 Engine HMS1700 includes a valve 1704a' positioned within air line 1702. Valve 1704a' is configured to direct air through heat exchanger 1703 or around heat exchanger 1703. Valve 1704a' is controlled by an actuator 1704' driven in a fuel hydraulic manner.

[0410] Actuator 1704' is also configured to actuate valve 1704a' to enable non-binary position adjustment between an open valve position and a closed valve position (i.e., actuator 1704' can simulate rather than binary control valve 1704a', or it can enable the valve to be adjusted to one or more intermediate positions between open and closed). Thus, valve 1704a' and associated fuel hydraulic actuator 1704' are configured to control the amount of air passing through heat exchanger 1803 to regulate the cooling of the oil.

[0411] In an alternative embodiment, Figure 22 and Figure 23 engines HMS can be combined such that HMS includes valves 1704a, 1704a' actuated via fuel hydraulic actuators 1704, 1704' in both the air line and the oil line.

[0412] In Figure 22 and Figure 23 the heat exchanger 1703 is a parallel flow heat exchanger. In other embodiments, the heat exchanger 1703 can be a counterflow heat exchanger, which means the flow direction of the air line or the oil line is reversed.

[0413] Figure 24Another embodiment is shown, in which the engine HMS1700 includes three heat exchangers 1703a - 1703c (located on branches of the oil line 1701) arranged parallel to the oil flow. Although Figure 24 the embodiment includes three heat exchangers, in other embodiments, any number of heat exchangers may be used. In the illustrated embodiment, all three heat exchangers 1703a - 1703c are air - oil heat exchangers.

[0414] The engine HMS1700 includes an oil line 1701 that supplies oil to the heat exchangers 1703a - 1703c. The engine HMS1700 includes three air lines 1702a - 1702c that respectively supply air to the heat exchangers 1703a - 1703c.

[0415] The engine HMS1700 includes a valve 1704a in the oil line 1701, which is configured such that oil can travel through or bypass all of the heat exchangers 1703a - 1703c. The valve 1704a is controlled by an actuator 1704 driven in a fuel - hydraulic manner.

[0416] The actuator 1704 is configured to actuate the valve 1704a so as to enable non - binary position adjustment (i.e., the valve can be continuously adjustable or adjustable between three or more positions) between an open - valve position and a closed - valve position. Thus, the valve 1704a and the associated fuel - hydraulic actuator 1704 are configured to control the amount of oil passing through the heat exchangers 1703a - 1703c in order to regulate the cooling of the oil.

[0417] In other embodiments, the engine HMS1700 includes additional valves, each valve being disposed within the air lines 1702a - 1702c and configured such that air can bypass the corresponding air - oil heat exchanger 1703a - 1703c. One or more of the valves disposed within the air lines 1902a - 1902c can be driven in a fuel - hydraulic manner.

[0418] Figure 25 An alternative embodiment is shown, in which the engine HMS1700 also includes three heat exchangers 1703a - 1703c arranged in parallel, but in which the valve arrangement is different from Figure 24 the valve arrangement of Figure 25 the embodiment includes three heat exchangers, in other embodiments, any number of heat exchangers may be used.

[0419] The engine HMS1700 includes an oil line 1701 that supplies oil to the heat exchangers 1703a - 1703c. The engine HMS1700 includes three air lines 1702a - 1702c that respectively supply air to the heat exchangers 1703a - 1703c.

[0420] The engine HMS1700 includes three valves 1704a, 1704a', 1704a'' (collectively 1704a), each valve 1704a being positioned in the oil line 1701, each valve being located on a different branch thereof. Each valve 1704a is configured such that oil can travel through or bypass the associated heat exchangers 1703a - 1703c. The valve 1704a enables oil to bypass one or more of the heat exchangers 1703a - 1703c in order to control the rate and amount of heat transferred from the oil. Each of the valves 1704a is controlled via a respective actuator 1704, 1704', 1704'' (collectively 1704). One or more of the actuators 1704 can be driven by fuel hydraulics.

[0421] The actuator 1704 is configured to actuate the valve 1704a such that non - binary position adjustment (i.e., the valve can be continuously adjustable or adjustable between three or more positions) can be made between the open - valve position and the closed - valve position of each valve 1704a. Thus, the valve 1704a and the associated fuel - hydraulic actuator 1704 are configured to individually control the amount of oil passing through each of the heat exchangers 1703a - 1703c in order to regulate the cooling of the oil.

[0422] In other embodiments, the engine HMS1700 includes additional valves, each valve being disposed within the air lines 1702a - 1702c and configured such that air can bypass the corresponding air - oil heat exchangers 1703a - 1703c. One or more of the valves disposed within the air lines 1702a - 1702c can be driven by fuel hydraulics.

[0423] In other embodiments, Figure 24 and Figure 25 the engine HMS systems 1700 can be combined such that the HMS system includes valves capable of bypassing all heat exchangers simultaneously and valves for bypassing individual heat exchangers.

[0424] Figure 26 A schematic diagram of the engine HMS1700 including a fuel - oil heat exchanger 1703a and an air - oil heat exchanger 1703b is shown. The engine HMS1700 includes an oil line 1701 having a valve 1704a. The valve 1704a can be operated by an actuator 1704 driven by fuel hydraulics. The valve 1704a divides the oil line 1701 into three sections or branches and regulates the flow of oil into each section.

[0425] A first portion 1701a of the oil pipeline passes through the fuel - oil heat exchanger 1703a. Fuel passes through the fuel - oil heat exchanger 1703a via the fuel pipeline 1705. The fuel in the fuel pipeline 1705 passing through the fuel - oil heat exchanger 1703a can be used to hydraulically drive the actuator 1704 in a fuel - hydraulic manner to control the valve 1704a. After the fuel has passed through the heat exchanger 1703a, or before the fuel passes through the heat exchanger 1703a, the fuel can be used to hydraulically drive the actuator 1704 in a fuel - hydraulic manner to control the valve 1704a.

[0426] A second portion of the oil pipeline 1701b does not pass through any of the heat exchangers 1703a, 1703b.

[0427] A third portion of the oil pipeline 1701c passes through the air - oil heat exchanger 1703b. Air passes through the air - oil heat exchanger 1703b via the air pipeline 1702.

[0428] Figure 26 The engine HMS1700 may include one or more additional valves in the oil pipeline 1701c and / or the oil pipeline 1701a. These valves can be hydraulically driven by fuel. The valves can enable the oil to bypass the associated heat exchanger. In such embodiments, the oil pipeline 1701b can be omitted. In other embodiments, one or more additional valves can be included in the air pipeline 1702 and / or the fuel pipeline 1705 so that air or fuel (if applicable) can bypass the respective heat exchangers 1703a, 1703b.

[0429] Figure 27 A schematic diagram showing a portion of the fuel supply system 152 is presented. The fuel supply system 152 includes a valve 1804a that is configured to split the fuel into two paths. The actuator 1804 hydraulically driven by fuel can be used to operate the valve 1804a. The first path 1801 diverts the fuel upstream of the fuel - oil heat exchanger 1803 such that the fuel travels through the fuel - oil heat exchanger 1803 on its way to the burner 16. The second path 1802 diverts the fuel downstream of the fuel - oil heat exchanger 1803 such that the fuel is not heated on its way to the burner 16. This portion of the fuel supply system 152 can be used to return the fuel to the main fuel flow path 110. The valve 1804a can be used to determine where along the fuel flow path and how much fuel is diverted. The fuel - oil heat exchanger 1803 can be Figure 26 the fuel - oil heat exchanger 1703a of the embodiment, or a different fuel - oil heat exchanger.

[0430] It should be understood that fuel for fuel hydraulic actuation from one or more fuel flow circuits 159, 259, etc. can return to the main fuel flow path 110 at any suitable point. In some cases, the fuel can even return to the fuel tanks 50, 53, but more commonly the fuel is kept within the engine 10 and simply recirculated within the engine 10 until it is sent to the burner 16 for combustion. The ability to control the return location of the fuel along the main fuel flow path 110 can contribute to overall engine thermal management. For example, if the fuel temperature of the fuel leaving one or more actuators is below a threshold, it can return to a point along the flow path 110 before the fuel-oil heat exchanger to allow the fuel temperature to further increase before it reaches the burner 16. Conversely, if the fuel temperature of the fuel leaving one or more actuators is above a threshold, it can return to a point along the flow path 110 after the fuel-oil heat exchanger, as it may not be desirable to further input heat to the fuel, and also optionally return to a point downstream of the engine pump to reduce the risk of the pump suffering thermal damage.

[0431] In an embodiment having multiple heat exchangers along the main fuel flow path 110, the valve 1804a can be configured to direct the fuel through or around one or more of the heat exchangers.

[0432] The generator HMS 1900 can be independent of the engine HMS 1700. It provides cooling for an electric machine, such as an electric machine used to supply power to the airframe to operate aircraft systems. The generator HMS 1900 includes one or more generator HMS valves 1904a, 1904a' controlled by one or more actuators 1904, 1904'. In various embodiments, one or more of these actuators 1904, 1904' are hydraulically driven by fuel.

[0433] For example, in some embodiments, the generator HMS 1900 includes a single valve and a single actuator hydraulically driven by fuel.

[0434] The generator HMS 1900 system can include multiple heat exchangers arranged in a parallel configuration. A single valve controlled by a single fuel hydraulic actuator can enable the use (or bypass) of the heat exchangers. In some embodiments, each heat exchanger has an associated valve controlled by an associated actuator. In various embodiments, one or more of these actuators are hydraulically driven by fuel.

[0435] Figure 28A schematic diagram showing a part of the generator HMS1900 is presented. The generator HMS1900 includes an oil-air heat exchanger 1903 which is configured to transfer heat from the oil within the engine, and more specifically from the oil used for cooling (and optionally also lubricating the generator). The generator HMS1900 includes an oil pipeline 1901 and an air pipeline 1902.

[0436] The generator HMS1900 includes a valve 1904a positioned within the oil pipeline 1901. The valve 1904a is configured to direct the oil through the heat exchanger 1903 or around the heat exchanger 1903. The valve 1904a is controlled by an actuator 1904 driven in a fuel hydraulic manner.

[0437] The actuator 1904 is configured to actuate the valve 1904a so as to enable continuous position adjustment between an open valve position and a closed valve position, thereby continuously regulating the oil flow rate through the air-oil heat exchanger 1903. Thus, the valve 1904a and the associated fuel hydraulic actuator 1904 are configured to control the amount of oil passing through the heat exchanger 1903 in order to regulate the cooling of the oil.

[0438] In other embodiments, the oil-air heat exchanger 1902 may be replaced by, or used in addition to, a fuel-oil heat exchanger. In such embodiments, fuel may be used to drive the actuator 1904 in a fuel hydraulic manner before the fuel is heated by the oil in the heat exchanger. Alternatively, fuel may be used to drive the actuator 1904 in a fuel hydraulic manner after the fuel has been heated by the oil in the heat exchanger.

[0439] In various embodiments, in addition to or instead of a valve provided in the oil pipeline, a valve and an associated actuator controllable in a fuel hydraulic manner may be provided in the air pipeline (for the oil-air heat exchanger) or in the fuel pipeline (for the fuel-oil heat exchanger).

[0440] Figure 29 A schematic diagram showing a part of the generator HMS1900 according to one such alternative embodiment is presented. The generator HMS1900 includes an oil-air heat exchanger 1903 which is configured to transfer heat from the oil within the engine, and more specifically from the oil used for cooling (and optionally also lubricating the generator). The generator HMS1900 includes an oil pipeline 1901 and an air pipeline 1902.

[0441] The generator HMS1900 includes a valve 1904a' positioned within an air line 1902. The valve 1904a' is configured to direct air through a heat exchanger 1903 or to bypass the heat exchanger 1903. The valve 1904a' is controlled by an actuator 1904' driven in a fuel hydraulic manner.

[0442] The actuator 1904' is configured to actuate the valve 1904a' so as to enable continuous position adjustment between an open valve position and a closed valve position. Accordingly, the valve 1904a' and the associated fuel hydraulic actuator 1904' are configured to control the amount of air passing through the heat exchanger 1903 in order to regulate the cooling of the oil.

[0443] In an alternative embodiment, Figure 28 and Figure 29 the HMS generators may be combined such that the generator HMS1900 includes valves actuated via fuel hydraulic actuators in both the air line and the oil line.

[0444] In Figure 28 and Figure 29 the heat exchanger 1903 is a parallel flow heat exchanger. In other embodiments, the heat exchanger 1903 may be a counterflow heat exchanger, which means that the flow direction of the air line or the oil line is reversed.

[0445] Figure 30 A schematic diagram showing a portion of the combined HMS2000 is presented. The combined HMS2000 includes an oil-oil heat exchanger 2005 configured to enable heat transfer between the oil in the generator HMS1900 (which for brevity may be referred to as generator oil) and the oil in the engine HMS1700 (which for brevity may be referred to as main engine oil). Accordingly, heat may be transferred between the two systems, effectively combining them into a single combined thermal management system while maintaining oil fluid isolation.

[0446] In Figure 30 the heat exchanger 2005 is a parallel flow heat exchanger. In other embodiments, the flow direction of either fluid passing through the heat exchanger may be reversed, and the heat exchanger 2005 may thus be a counterflow heat exchanger.

[0447] The main engine oil travels via the oil line 1701 of the engine HMS1700 into the heat exchanger 2005. In some embodiments, the oil line 1701 includes a valve 1704 configured to divert some or all of the oil to bypass the heat exchanger 2005. The valve 1704 may be operated using an actuator 1704' driven in a fuel hydraulic manner.

[0448] The generator oil travels through the oil line 1901 of the generator HMS1900 into the heat exchanger 2005. In some embodiments, the oil line 1901 includes a valve 1904a”, which is configured to divert some or all of the oil to bypass the heat exchanger 2005. The valve 1904a” can be operated using an actuator 1904” driven in a fuel hydraulic manner.

[0449] Various methods for operating a gas turbine engine of an aircraft are described below.

[0450] Figure 31 A flowchart of a method 1000 for operating a gas turbine engine 10 according to various embodiments is shown. The method 1000 is performed on the engine 10, which includes: an engine core 11, which includes a turbine 19, a combustor 16, a compressor 14, and a spool 26 connecting the turbine to the compressor; a fan 23, which is located upstream of the engine core 11 and is arranged to be driven by the spool 26, the fan including a plurality of fan blades; and a nacelle 21, which surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 located radially outside the engine core 11; a plurality of actuators 254; and a fuel supply system 152.

[0451] In the described embodiments, the bypass ratio is greater than or equal to 4, and the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct 22 to the mass flow rate of the flow through the core 11 under cruise conditions. Such a bypass ratio can be used for any engine performing Figures 31 to 35 any method.

[0452] Method 1000 includes:

[0453] Supplying 1010 fuel using the fuel supply system 152 for combustion in the combustor 16;

[0454] And

[0455] Supplying 1020 fuel using the fuel supply system 152 to hydraulically drive at least ten of the plurality of actuators 254 by fuel.

[0456] The at least ten actuators 254 can include:

[0457] · Any one of at least two valves 554a, 554b within the VSV system 556;

[0458] · Any one of at least two valves 654a, 654a’ within the TCC system 656;

[0459] · Any one of at least three valves 754a, 754a’, 754a” within the HMU 756;

[0460] · A ventilation valve actuator 854;

[0461] · The nacelle anti-ice valve actuator 954 or any other anti-ice valve actuator;

[0462] · The bleed air actuator 1054;

[0463] · The regulating component actuator 1154;

[0464] · The advanced valve actuator 1254;

[0465] · The actuator 1354 for controlling the valve that extracts bleed air from non-compressor components of the engine;

[0466] · The APU bleed air actuator 1454;

[0467] · The engine start valve actuator 1554;

[0468] · The isolation valve actuator 1654;

[0469] · Any one of the actuators 1704, 1704' of the engine thermal management system 1700;

[0470] and / or

[0471] · Any one of the actuators 1904, 1904' of the generator thermal management system 1900.

[0472] Then, fuel for hydraulically driving one or more of the actuators 254 in a fuel-powered manner can be supplied to the burner 16 for combustion. The fuel can flow along the main fuel flow path 110 to the burner 16, so the main fuel flow path 110 can supply fuel 1010 to the burner 16. A portion of the fuel flowing along the main fuel flow path 110 can be diverted from the main fuel flow path 110 to the actuator 254, and then the fuel can return to the main fuel flow path 110 after being used for actuation. One or more tubes 159, 160 can be used to convey fuel from the main fuel flow path 110 to the actuator 254. These tubes can form one or more fuel flow circuits. One or more valves can be used to control the fuel flow rate through the or each fuel hydraulic flow circuit. In some embodiments, one or more fuel hydraulic fuel pumps can be provided to actively pump fuel to or from the actuator 254. However, in many embodiments, the pressure provided by one or more fuel pumps on the main fuel flow path 110 may be sufficient.

[0473] The discharge device 159 on the main fuel flow path 110 can be located at a point different from the return pipe 160 of (or each) fuel flow circuit along the main fuel flow path 110. The fuel can return to the main fuel flow path 110 at a position upstream or downstream of this discharge device. One or more valves can be used to control the position along the main fuel flow path 110 at which the fuel for driving the actuator returns to the main fuel flow path 110. For example, the fuel for actuation can return to the fuel tanks 50, 53, return just before the burner 16, or return before or after any fuel - oil heat exchanger or fuel pump, as the case may be. Generally, the fuel for actuation returns to a point on the main fuel flow path 110 within the engine 10 (rather than, for example, directly returning to the fuel tanks 50, 53 located at other positions in the aircraft 1 or returning to the connecting pipe between them).

[0474] Accordingly, the supply steps 1010, 1020 can include controlling a plurality of fuel flow valves and optionally also include controlling a plurality of fuel pumps.

[0475] When actuation is not required, the percentage of fuel redirected to the actuator can be 0%.

[0476] The fuel flow rate to the actuator under idle conditions is typically equal to or even higher than the flow rate to the burner 16 under idle conditions, for example, between one time and 5.5 times the fuel flow rate to the burner 16.

[0477] The flow rate to the actuator under cruise conditions is typically lower than or similar to the flow rate to the burner 16, for example, between 0.26 times and 1.1 times the fuel flow rate to the burner 16.

[0478] Figure 32 A flowchart of a method 1100 for operating a gas turbine engine 10 according to various embodiments is shown. The method 1100 is performed on the engine 10, which includes: an engine core 11 including a turbine 19, a burner 16, a compressor 14, and a spool 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and arranged to be driven by the spool 26, the fan including a plurality of fan blades; a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuated engine systems 356, 456, 556, 656, 756 including a thermal management system 1700, 1900, and a turbine housing cooling system 656; and a fuel supply system 152.

[0479] The method 1100 includes:

[0480] Supplying 1110 fuel using the fuel supply system for combustion in the burner; and

[0481] Supply fuel to drive at least three of a plurality of actuated engine systems in a fuel hydraulic manner using a fuel supply system.

[0482] The at least three actuated engine systems may include one or more of the following:

[0483] ·VSV system 556;

[0484] ·TCC system 656;

[0485] ·HMU 756;

[0486] ·Engine thermal management system 1700;

[0487] ·Generator thermal management system 1900; and / or

[0488] ·Any one of the bleed air systems in the bleed air system.

[0489] During cruise conditions, the fuel flow rate to each actuator within the thermal management system can be 0% to 15% of the fuel flow rate to burner 16.

[0490] During idle conditions, the fuel flow rate to each actuator within the thermal management system can be 0% to 150% of the fuel flow rate to burner 16.

[0491] In some embodiments, there can be between 1 and 5 fuel hydraulic actuators in the thermal management system.

[0492] During cruise conditions, the fuel flow rate to the TCC system can be 0% to 30% of the fuel flow rate to burner 16, and optionally in the range of 0% to 20%.

[0493] During idle conditions, the fuel flow rate to the TCC system can be 0% to 150% of the fuel flow rate to burner 16, and optionally in the range of 0% to 75%.

[0494] As described for Figure 31 method 1000, one or more fuel flow circuits and associated valves can be used to control how much fuel is diverted from the main fuel flow path 110 to the actuated systems, and where the fuel returns along the main fuel flow path 110.

[0495] Figure 33 A flowchart of a method 1200 for operating a gas turbine engine according to various embodiments is shown. Method 1200 is performed on engine 10, which includes:

[0496] The engine core 11, which includes a turbine 19, a combustor 16, a compressor 14, and a spool 26 that connects the turbine to the compressor; a fan 23, which is located upstream of the engine core 11 and is arranged to be driven by the spool 26, the fan including a plurality of fan blades; a nacelle 21, which surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 located radially outside the engine core 11; a plurality of actuators (any one or all of actuators 954, 1054, 1154, 1254, which may be more generally represented as 254), wherein at least one of the plurality of actuators is configured to actuate bleed valves 954a, 1054a, 1154a, 1254a; and a fuel supply system 152.

[0497] The method 1200 includes:

[0498] Using the fuel supply system 152 to supply 1210 fuel for combustion in the combustor 16;

[0499] And

[0500] Using the fuel supply system 152 to supply 1220 fuel to hydraulically drive at least one of the actuators 954, 1054, 1154, 1254 that is configured to actuate the bleed valves 954a, 1054a, 1154a, 1254.

[0501] The bleed valve can be any other bleed valve as described above. For example, the bleed valve can be an anti-icing valve. The bleed valve can be an advanced valve. The bleed valve can be a manifold pressure valve. The bleed valve can be a high-pressure bleed valve, a medium-pressure valve, or a bypass duct.

[0502] As described with respect to Figure 31 And Figure 32 The methods 1000, 1100, one or more fuel flow circuits and associated valves can be used to control how much fuel is diverted from the main fuel flow path 110 to the actuated system, and where the fuel returns along the main fuel flow path 110.

[0503] Figure 34FIG. 1300 is a flow chart showing a method 1300 of operating a gas turbine engine according to various embodiments. Method 1300 is performed on an engine 10 that includes: an engine core 11 that includes a turbine 19, a combustor 16, a compressor 14, and a spool 26 that connects the turbine to the compressor; a fan 23 that is located upstream of the engine core 11 and is arranged to be driven by the spool 26, the fan including a plurality of fan blades; a nacelle 21 that surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 that is located radially outward of the engine core 11; an engine thermal management system 1700; a plurality of actuators, including actuators 1704, 1704' that are configured to actuate valves 1704a, 1704a' within the engine thermal management system 1700, wherein the actuators 1704, 1704' that are configured to actuate the valves 1704a, 1704a' within the engine thermal management system 1700 are configured to actuate the valves 1704a, 1704a' such that non-binary position adjustment between an open valve position and a closed valve position is possible; and a fuel supply system 152.

[0504] Method 1300 includes:

[0505] Supplying 1310 fuel using the fuel supply system 152 for combustion in the combustor 16;

[0506] And

[0507] Supplying 1320 fuel using the fuel supply system 152 to hydraulically drive at least the engine thermal management system valve actuators 1704', 1704a' among the plurality of actuators.

[0508] Engine 10 of some embodiments includes a plurality of actuators 1704, 1704', each actuator being configured to actuate valves 1704a, 1704a' within the engine HMS 1700, and method 1300 of such embodiments may include hydraulically actuating the plurality of such actuators using the fuel supply system 152.

[0509] The engine HMS system 1700 may control the temperature of an oil system that lubricates engine components (e.g., bearing housings, gearboxes, etc.). The engine HMS system 1700 includes one or more engine HMS valves 1704a, 1704a' that are controlled by one or more fuel hydraulic actuators 1704, 1704'.

[0510] As described with respect to Figures 31 to 33 methods 1000, 1100, 1200, one or more fuel flow circuits and associated valves may be used to control how much fuel diverges from the main fuel flow path 110 to an actuated system, and where the fuel returns along the main fuel flow path 110.

[0511] Figure 35 FIG. 1400 shows a flow chart of a method 1400 of operating a gas turbine engine according to various embodiments. The method 1400 is performed on an engine 10 that includes: an engine core 11 that includes a turbine 19, a combustor 16, a compressor 14, and a shaft 26 that connects the turbine to the compressor; a fan 23 that is located upstream of the engine core 11 and is arranged to be driven by the shaft 26, the fan including a plurality of fan blades; a nacelle 21 that surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 that is radially outside of the engine core 11; a generator heat management system 1900; a plurality of actuators, including actuators 1904, 1904' that are configured to actuate valves 1904a, 1904a' within the generator heat management system 1900, wherein the actuators 1904', 1904a' that are configured to actuate the valves 1904a, 1904a' within the generator heat management system 1900 are configured to actuate the valves 1904a, 1904a' such that a non-binary position adjustment between an open valve position and a closed valve position is enabled; and a fuel supply system 152.

[0512] The method 1400 includes:

[0513] supplying 1410 fuel using the fuel supply system 152 for combustion in the combustor 16;

[0514] and

[0515] supplying 1420 fuel using the fuel supply system 152 to hydraulically drive at least the generator heat management system valve actuators 1904, 1904' of the plurality of actuators by fuel.

[0516] The engine 10 of some embodiments includes a plurality of actuators 1904, 1904', each actuator being configured to actuate valves 1904a, 1904a' within the generator HMS 1900, and the method 1400 of such embodiments may include hydraulically actuating the plurality of such actuators using the fuel supply system 152 by fuel.

[0517] The generator HMS 1900 may be independent of the engine HMS 1700. It may provide cooling for an electric machine, such as for supplying power to a fuselage to operate aircraft systems. The generator HMS 1900 includes one or more generator HMS valves 1904a, 1904' that are controlled by one or more fuel hydraulic actuators 1904, 1904'.

[0518] As regarding Figures 31 to 34As described in methods 1000, 1100, 1200, 1300, one or more fuel flow circuits and associated valves can be used to control how much fuel diverges from the main fuel flow path 110 to the actuated system, and where the fuel returns along the main fuel flow path 110.

[0519] It should be understood that the present invention is not limited to the above-described embodiments, 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, the gas turbine engine comprising: an engine core, the engine core comprising a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor; a fan located upstream of the engine core and arranged to be driven by the spindle, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct radially outward of the engine core, wherein a bypass ratio is at least 4, the bypass ratio being defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core under cruise conditions; multiple actuators; and A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least ten actuators of the plurality of actuators via the fuel.

2. The gas turbine engine of claim 1, wherein the fuel supply system is arranged to supply fuel to hydraulically drive at least eleven actuators of the plurality of actuators via the fuel.

3. The gas turbine engine of claim 2, wherein the fuel supply system is arranged to supply fuel to hydraulically drive at least twelve of the plurality of actuators via the fuel.

4. A gas turbine engine according to claim 1, wherein the engine includes a variable stator vane system, wherein at least two of the plurality of actuators are part of the variable stator vane system, and wherein the fuel supply system is arranged to supply fuel to hydraulically drive the at least two of the plurality of actuators that are part of the variable stator vane system.

5. A gas turbine engine according to claim 1, wherein the engine includes a turbine casing cooling system, wherein at least two of the plurality of actuators are part of the turbine casing cooling system, and wherein the fuel supply system is arranged to supply fuel to hydraulically drive the at least two of the plurality of actuators that are part of the turbine casing cooling system.

6. The gas turbine engine of claim 1, wherein the engine comprises a ventilation valve actuator, and wherein the fuel supply system is arranged to supply fuel to hydraulically drive the ventilation valve actuator via fuel.

7. A gas turbine engine according to claim 1, wherein the engine includes an engine thermal management system, and wherein at least one of the plurality of actuators is part of the engine thermal management system, and wherein the fuel supply system is arranged to supply fuel to drive at least one of the plurality of actuators as part of the engine thermal management system by fuel hydraulic means.

8. A gas turbine engine according to claim 1, wherein the engine includes a generator thermal management system, and wherein at least one actuator of the plurality of actuators is part of the generator thermal management system, and wherein the fuel supply system is arranged to supply fuel to drive the at least one actuator of the plurality of actuators that is part of the generator thermal management system by fuel hydraulic means.

9. A gas turbine engine according to claim 1, wherein the engine includes a hydraulic-mechanical unit, and wherein at least three of the plurality of actuators are part of the hydraulic-mechanical unit, and wherein the fuel supply system is arranged to supply fuel to drive the at least three of the plurality of actuators that are part of the hydraulic-mechanical unit by fuel hydraulic means.

10. A gas turbine engine according to claim 9, wherein at least four of the plurality of actuators are part of the hydraulic-mechanical unit, and wherein the fuel supply system is arranged to supply fuel to hydraulically drive the at least four of the plurality of actuators that are part of the hydraulic-mechanical unit by fuel.

11. A gas turbine engine according to claim 1, wherein the engine comprises an induction valve, wherein at least one actuator of the plurality of actuators is configured to actuate the induction valve, and wherein the fuel supply system is arranged to supply fuel to hydraulically drive the at least one actuator of the plurality of actuators configured to actuate the induction valve via fuel.

12. The gas turbine engine of claim 1, wherein the spindle outputs drive directly to the fan, thereby driving the fan at the same rotational speed as the spindle, such that the engine is a direct drive turbine engine.

13. The gas turbine engine of claim 1, wherein the turbine engine includes a gearbox receiving input from the spindle and outputting drive to the fan, thereby driving the fan at a lower rotational speed than the spindle, such that the engine is a geared turbine engine.

14. A gas turbine engine according to claim 1, wherein the engine includes a plurality of actuated engine systems, including a thermal management system and a turbine casing cooling system, and wherein the fuel supply system is arranged to supply fuel to drive at least three of the plurality of actuated engine systems by fuel hydraulic means.

15. The gas turbine engine of claim 14, wherein the fuel supply system is arranged to supply fuel to hydraulically drive at least five of the plurality of actuated engine systems via fuel.

16. The gas turbine engine of claim 15, wherein the fuel supply system is arranged to supply fuel to hydraulically drive at least seven of the plurality of actuated engine systems via fuel.

17. A gas turbine engine according to claim 14, wherein at least one of the multiple actuated engine systems includes at least two of the multiple actuators, and wherein the fuel supply system is arranged to supply fuel to drive at least one of the at least two of the multiple actuators within the at least one actuated engine system by fuel hydraulic means.

18. A gas turbine engine according to claim 14, wherein at least one of the actuated engine systems includes at least two of the plurality of actuators, and wherein the fuel supply system is arranged to supply fuel to drive each of the at least two of the plurality of actuators within the at least one actuated engine system by fuel hydraulic means.

19. A method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core, the engine core comprising a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor; a fan located upstream of the engine core and arranged to be driven by the spindle, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass radially outward of the engine core, wherein a bypass ratio is at least 4, the bypass ratio being defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core at cruise conditions; multiple actuators; and Fuel supply system; And wherein the method comprises: supplying fuel for combustion in the burner using the fuel supply system; as well as Fuel is supplied using the fuel supply system to hydraulically drive at least ten actuators of the plurality of actuators with the fuel.

20. The method of claim 19, comprising supplying fuel to hydraulically drive at least eleven actuators of the plurality of actuators with the fuel.

Images