Fuel-driven hydraulic actuation

Abstract

Fuel-driven hydraulic actuation A gas turbine engine for an aircraft includes an engine core including a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and configured to be driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outward of the engine core, where the bypass ratio, 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 at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators and a fuel delivery system.The fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel for hydraulically fuel-driving at least ten of the plurality of actuators. Figure for abstract: Fig. 31.

Description

Title of the invention: Fuel-driven hydraulic actuation

[0001] The present description relates to aircraft actuation systems, and to methods of controlling a hydraulically fueled actuation system such that one or more actuators of the system are hydraulically fueled.

[0002] There is an expectation in the aviation industry of a trend toward the use of fuels different from the traditional kerosene-based jet fuels generally used today. The fuels may have different fuel characteristics than petroleum-based hydrocarbon fuels. Thus, there is a need to consider fuel properties for these new fuels, and to adjust both the gas turbine engines themselves and the methods of operating the gas turbine engines.

[0003] According to a first aspect, there is provided a gas turbine engine for an aircraft, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators; and a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least ten of the plurality of actuators.

[0004] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that these different fuel properties may allow the actuators to be hydraulically driven by fuel. In particular, certain fuels can be heated to higher temperatures and used to actuate at least one more actuator than traditional fuels, without significantly increasing significantly reduces the risk of thermal degradation of the fuel (e.g., fuel lacquer, or fuel coking) in the actuators.

[0005] 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.

[0006] The fuel supply system may include a fuel tank return valve. The fuel tank return valve may be controlled using a fuel tank return actuator. The fuel supply system may be configured to supply fuel to the fuel tank return actuator to hydraulically drive it with fuel.

[0007] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least eleven, or at least twelve of the plurality of actuators.

[0008] The engine may include a variable vane stator system. At least two of the plurality of actuators may be part of the variable vane stator system and may actuate / move the vanes in use. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least two of the plurality of actuators that are part of the variable vane stator system.

[0009] The engine may include a turbine case cooling, TCC, system. At least two of the plurality of actuators may be part of the turbine case cooling system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least two of the plurality of actuators that are part of the TCC system.

[0010] The engine may include a ventilation valve actuator. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the ventilation valve actuator.

[0011] The engine may include an engine heat management system. The engine heat management system may be configured to cool oil used to cool and lubricate one or more engine components, including shaft bearings and a reduction gear (if applicable). The engine heat management system may be configured to control the temperature of the fuel at the inlet to the combustion chamber. The engine heat management system may include a plurality of heat exchangers; for example, one or more air-to-oil heat exchangers and one or more fuel-to-oil heat exchangers. At least one of the plurality of actuators may be part of the engine heat management system. The fuel supply system is configured to supply fuel to hydraulically fuel-drive the at least one of the plurality of actuators in the engine heat management system. Various actuators in the engine heat management system may be used to control fluid flow (e.g., air, oil, or fuel flow) through one or more heat exchangers in the engine heat management system.

[0012] The engine may include a generator heat management system. At least one of the plurality of actuators is part of the generator heat management system. The fuel supply system is configured to supply fuel to hydraulically fuel-drive the at least one of the plurality of actuators in the generator heat management system.

[0013] The engine may include a hydromechanical unit. At least three of the plurality of actuators may be part of the hydromechanical unit. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least three of the plurality of actuators that are part of the hydromechanical unit.

[0014] At least four of the plurality of actuators may be part of the hydromechanical unit. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least four of the plurality of actuators that are part of the hydromechanical unit.

[0015] The engine may include an air bleed valve. At least one of the plurality of actuators may be configured to actuate the air bleed valve. The fuel delivery system may be configured to provide fuel to hydraulically fuel-drive at least one of the plurality of actuators configured to actuate the air bleed valve.

[0016] The at least ten fuel hydraulic actuators (i.e., actuators that are hydraulically driven by fuel) may therefore comprise one or more of the following: • at least one variable stator vane actuator, VSV, designed to adjust the position of a stator vane of the motor; • at least one actuator designed to actuate a valve in a turbine case cooling, TCC, system of the engine; • at least one actuator designed to operate a valve in a hydromechanical unit, HMU, of the engine; • at least one actuator designed to operate a ventilation valve; • at least one actuator designed to operate a nacelle anti-icing valve or any other anti-icing or defrosting valve; • at least one actuator designed to operate an air bleed valve; • at least one actuator designed to operate a conditioning valve; • at least one actuator adapted to operate an upper stage valve to control an air flow from the compressor; • at least one actuator designed to operate a valve for the extraction of bleed air from parts of the engine which are not compressors; • at least one actuator configured to actuate a bleed air valve of an auxiliary power unit, APU (in some embodiments, however, the APU may be separate from the engine and an APU actuator may therefore not be considered part of the engine. Fuel from a portion of the fuel supply path upstream of the engine may be used to actuate such an APU actuator. In other embodiments, however, the APU may be associated with or part of the engine, and the AP actuator may in such cases be considered part of the engine); • an engine start valve actuator; • an isolation valve actuator; • at least one actuator designed to operate a valve of an engine heat management system; and / or • at least one actuator designed to operate a valve of a generator heat management system.

[0017] For example, an engine heat management system typically includes multiple servomotors (also referred to as servos), a servo being a rotary or linear actuator that provides precise control of angular or linear position, velocity, and / or acceleration in a mechanical system. In a given engine heat management system, one, more, or all of the servos present may be hydraulically fuel-actuated. One or more of the servomotors may include a sensor configured to provide position feedback. A specific controller may be provided for the engine heat management system to control the one or more servomotors. The same is true for the generator heat management system.

[0018] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, so that the engine is a direct-drive turbine engine.

[0019] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0020] The engine may include a plurality of powered engine systems. The powered engine systems may include one or more of: a heat management system, a turbine case cooling system, a generator heat management system, an engine heat management system, a bleed air system, a variable vane stator system, VSV, a cabin environmental control system, a defrost system, and a bearing chamber venting system.

[0021] The fuel supply system may be configured to provide fuel to hydraulically fuel-drive at least three of the plurality of actuated engine systems. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive at least four, five, six, or at least seven of the actuated engine systems.

[0022] It will be appreciated that each actuated engine system may include multiple actuators, and that not all of the actuators in a given system may be fuel hydraulic. Therefore, fuel hydraulic driving of a given system includes actuation of at least one, but not necessarily all, of the actuators in the system.

[0023] 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 configured to provide fuel to hydraulically fuel-drive at least one of the at least two of the plurality of actuators in the at least one actuated engine system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive each of the at least two of the plurality of actuators in the at least one actuated engine system.

[0024] According to a second aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators; and a fuel supply system; the method comprising: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel-drive at least ten of the plurality of actuators.

[0025] The method of the second aspect may be performed using the engine of the first aspect.

[0026] According to a third aspect, a gas turbine engine for an aircraft is provided, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of powered engine systems, including a heat management system and a turbine case cooling system; and a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least three of the plurality of actuated engine systems.

[0027] Fuel hydraulic drive of an actuated engine system may mean fuel hydraulic drive of at least one actuator in the actuated engine system.

[0028] The powered engine systems may include one or more of: a heat management system, a turbine case cooling system, a generator heat management system, an engine heat management system, a bleed air system, a variable vane stator system, VSV, a cabin environmental control system, a defrost system, and a bearing chamber venting system.

[0029] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least four, five, six, or seven of the plurality of powered engine systems.

[0030] 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.

[0031] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the turbine case cooling system.

[0032] The engine may include an engine heat management system. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the engine heat management system.

[0033] The engine may include a generator heat management system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the generator heat management system.

[0034] The engine may include a ventilation valve system. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the ventilation valve system.

[0035] The engine may include an air bleed system. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the air bleed system.

[0036] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0037] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0038] The engine may include a plurality of actuators. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least ten, eleven, or twelve of the plurality of actuators.

[0039] At least one of the actuated engine systems may include at least two of the plurality of actuators. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive at least one of the at least two actuators in the at least one actuated engine system.

[0040] At least one of the actuated engine systems may include at least two of the plurality of actuators. The fuel supply system may be configured to provide fuel to hydraulically fuel each of the at least two actuators in the at least one actuated engine system.

[0041] The motor of the third aspect may be arranged to implement the method of the second aspect, and may have any of the characteristics described in relation to the first or second aspect.

[0042] According to a fourth aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and adapted to be driven by the core shaft (26), the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of powered engine systems, including a heat management system and a turbine housing cooling system; and a fuel delivery system; and wherein the method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel-drive at least three of the plurality of powered engine systems.

[0043] The method may include providing, using the fuel supply system, fuel to hydraulically fuel-drive at least four or five of the plurality of powered engine systems.

[0044] The engine may include an engine heat management system and a generator heat management system. The method may include providing fuel to hydraulically fuel-drive the engine heat management system and / or the generator heat management system.

[0045] The methods of the second and fourth aspects may be complementary, and may be performed together in various implementations. The method of the fourth aspect may be performed using the engine of the first aspect or the third aspect.

[0046] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators, at least one of the plurality of actuators being configured to actuate an air bleed valve; and a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive the at least one actuator configured to actuate the air bleed valve.

[0047] 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.

[0048] The bleed air flow rate may therefore be hydraulically controlled by fuel. The bleed air generally comprises air "bled" from the one or more compressors of the engine, and / or an auxiliary power unit (APU). The aircraft may include a bleed air system comprising a plurality of conduits, valves, and regulators configured to conduct this high-pressure air "bled" from the one or more compressors and / or the APU to different locations in the aircraft (both within the engine and elsewhere). The bleed air may be used for multiple different functions, including cabin pressurization, air conditioning, engine starting, bearing chamber venting, and deicing. At least a portion of the aircraft bleed air system may be part of, or contained within, the engine; this portion may be referred to as the engine bleed air system.The aircraft bleed air system may include multiple bleed air systems, some of which are located entirely within an engine.

[0049] The at least one air bleed valve may be an environmental air bleed valve. The bleed air may therefore be used to provide pressure to the aircraft cabin, providing air to an environmental control system. The at least one air bleed valve may be a conditioning unit valve. The conditioning valve (an example of an environmental air bleed valve) may regulate a bleed air flow rate in an air cycle air conditioning system.

[0050] Additionally or alternatively, the bleed air may be used to keep critical parts of the aircraft (such as the leading edges of the wings, parts of a nacelle surrounding the fan, and / or the engine section stator) free of ice. The bleed air valve may be referred to as an anti-icing valve.

[0051] The at least one air bleed valve may be a nacelle anti-icing valve.

[0052] The at least one air bleed valve may be a stator anti-icing valve. engine section.

[0053] The engine may include an isolation valve configured to allow inlet of bleed air from an external engine (i.e., the isolation valve may allow cross-feeding from the other / other engine of the same aircraft (it will be borne in mind that at least two engines (one per wing) are generally present for commercial aircraft, and many aircraft have multiple engines per wing, for example for a total of four or six engines. The bleed air may therefore be shared between the engines.). At least one of the plurality of actuators may be configured to actuate the isolation valve. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least one actuator configured to actuate the isolation valve.

[0054] 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 system may be configured to provide fuel to fuel-hydraulic drive the at least one actuator configured to actuate the engine starter valve. Bleed air, extracted from an auxiliary power unit (APU) or an operating engine of the aircraft, may therefore be used to power an air turbine starter motor to start an engine / other engine of the aircraft. The use of an air turbine starter may allow a smaller and lighter unit to provide the required torque compared to electric or hydraulic starters. Fuel-hydraulic control of the engine starter valve may further reduce the size and / or weight of the system.

[0055] The engine may include an upper stage valve for controlling a flow rate of air from the compressor. At least one of the plurality of actuators may be configured to actuate the upper stage valve. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least one actuator configured to actuate the upper stage valve. The upper stage valve may be configured to control the flow of high-pressure bleed air from the compressor to the aircraft cabin.

[0056] The engine may include, or be associated with, an auxiliary power unit, APU, and an auxiliary power unit valve. The auxiliary power unit may be separate from the engine, and mounted elsewhere on the aircraft in some implementations. At least one of the plurality of actuators may be configured to actuate the auxiliary power unit valve. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the at least one actuator configured to actuate the auxiliary power unit valve.

[0057] 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 system may be configured to provide fuel to hydraulically fuel-drive the 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 modulate / control the pressure of the bleed air flow from the compressor to the aircraft cabin.

[0058] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least ten, at least eleven, or at least twelve of the plurality of actuators.

[0059] The engine may include a plurality of powered engine systems. The plurality of powered engine systems may include a heat management system and a turbine housing cooling system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive at least three of the plurality of powered engine systems.

[0060] The powered engine systems may include one or more of: a heat management system, a turbine case cooling system, a generator heat management system, an engine heat management system, a bleed air system, a variable vane stator system, VSV, a cabin environmental control system, a defrost system, and a bearing chamber venting system.

[0061] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least four, five, six, or at least seven of the actuated engine systems.

[0062] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0063] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0064] The engine of the fifth aspect may be used to implement the method of the second and / or fourth aspect, and may have any of the characteristics described in relation to the first to fourth aspects.

[0065] According to a sixth aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators, at least one of the plurality of actuators being configured to actuate an air bleed valve; and a fuel delivery system; and wherein the method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel-drive the at least one actuator configured to actuate the air bleed valve.

[0066] The engine may include multiple air bleed valves. The method may include providing, using the fuel delivery system, fuel to hydraulically fuel-drive a plurality of actuators each configured to actuate an air bleed valve.

[0067] The methods of the second, fourth, and sixth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the sixth aspect may be performed using the engine of the first aspect, the third aspect, or the fifth aspect.

[0068] According to a seventh aspect, a gas turbine engine for an aircraft is provided comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and adapted to be driven by the core shaft (26), the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio, 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 at cruise conditions, is at least 4; an engine heat management system; a plurality of actuators, including an actuator configured to actuate at least one valve in the engine heat management system; a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive the actuator configured to actuate the at least one valve in the engine heat management system; wherein the actuator configured to actuate the at least one valve in the engine heat management system is configured to actuate said valve so as to allow a non-binary position adjustment between an open valve position and a closed valve position.

[0069] The actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate said valve in a manner that allows for position adjustment between an open valve position, one or more distinct intermediate valve positions, and a closed valve position. The non-binary position adjustment may therefore comprise a distinct position adjustment between at least three different valve positions. The actuator may be configured to actuate the valve between at least four defined positions. For example, the actuator may be configured to actuate the valve between four defined positions: open, closed, and two distinct / defined intermediate positions.

[0070] The actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate said valve so as to allow continuous position adjustment between an open valve position and a closed valve position.

[0071] 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.

[0072] The engine heat management system may include a plurality of valves.

[0073] The actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate at least two of the plurality of valves of the engine heat management system.

[0074] The plurality of actuators includes a plurality of actuators each configured to actuate at least one of the plurality of valves in the engine heat management system.

[0075] The engine may include an air bleed valve. The plurality of actuators may include an actuator configured to actuate the air bleed valve. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the actuator configured to actuate the air bleed valve.

[0076] The engine may include a generator heat management system. The plurality of actuators may include an actuator configured to actuate at least one valve in the generator heat management system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the actuator configured to actuate the at least one valve in the generator heat management system.

[0077] The generator heat management system may include a plurality of valves.

[0078] The at least one actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate at least two of the plurality of valves in the generator heat management system.

[0079] The plurality of actuators may include a plurality of actuators each configured to actuate at least one of the plurality of valves with the generator heat management system.

[0080] The actuator configured to actuate the at least one valve in the generator heat management system may be configured to actuate the at least one valve in the generator heat management system in a manner to allow continuous position adjustment between an open valve position and a closed valve position.

[0081] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least ten, eleven, or at least twelve of the plurality of actuators.

[0082] The engine may include a plurality of actuated engine systems. The engine systems may include a heat management system and a turbine case cooling system. The engine systems may include a VSV actuation system, one or more deicing systems, and / or a cabin air conditioning system. The fuel supply system may be configured to supply fuel to hydraulically fuel-driven at least three of the plurality of powered engine systems.

[0083] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least four, five, six, or seven of the powered engine systems.

[0084] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0085] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0086] The engine heat management system may include an air-to-oil heat exchanger. The actuator configured to actuate at least one valve in the engine heat management system may be configured to actuate an air-side valve of the air-to-oil heat exchanger. The actuator may therefore control air flow. The actuator configured to actuate at least one valve in the engine heat management system may be configured to actuate an air-side valve of the air-to-oil heat exchanger. The actuator may therefore control oil flow.

[0087] The engine heat management system may include one or more, and optionally a plurality of, air-to-oil heat exchangers.

[0088] The engine heat management system may include one or more, and optionally a plurality of, fuel-to-oil heat exchangers.

[0089] One or more of, and optionally each of, the plurality of heat exchangers may include a valve controllable via an associated actuator. The fuel supply system may be configured to supply fuel to hydraulically fuel each of the associated actuators.

[0090] The at least one valve in the engine heat management system may be configured to allow oil to bypass one or more of the plurality of air-to-oil heat exchangers.

[0091] The at least one valve in the engine heat management system may be configured to allow air to bypass one or more of the plurality of air-to-oil heat exchangers.

[0092] The at least one valve in the engine heat management system may be configured to allow fuel to bypass one or more of the plurality of fuel-to-oil heat exchangers.

[0093] The at least one valve in the engine heat management system may be configured to allow oil to bypass one or more of the plurality of fuel-to-oil heat exchangers.

[0094] The portion of fluid that can bypass a respective heat exchanger can be adjusted.

[0095] The motor of the seventh aspect may be configured to perform the method of the second, fourth, and / or sixth aspects, and may have any of the characteristics described in relation to the first to sixth aspects.

[0096] According to an eighth aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; an engine heat management system; a plurality of actuators, comprising an actuator configured to actuate a valve in the engine heat management system, wherein the actuator configured to actuate a valve in the engine heat management system is configured to actuate said valve so as to allow a non-binary position adjustment between an open valve position and a closed valve position; and a fuel supply system; and wherein the method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel-drive at least the actuator of the engine heat management system valve of the plurality of actuators.

[0097] The position adjustment may be continuous, or may be distinct between three or more different positions, and possibly between four or more different positions,

[0098] The methods of the second, fourth, sixth, and eighth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the eighth aspect may be performed using the engine of the first aspect, the third aspect, the fifth aspect, or the seventh aspect.

[0099] According to a ninth aspect, a gas turbine engine for an aircraft is provided, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a generator heat management system; a plurality of actuators, including an actuator configured to actuate at least one valve in the generator heat management system; and a fuel supply system, wherein the fuel supply system is configured to provide fuel for combustion in the combustion chamber, and to provide fuel to hydraulically fuel-drive the actuator configured to actuate the at least one valve in the generator heat management system; wherein the actuator configured to actuate the at least one valve in the generator heat management system is configured to actuate said valve so as to allow a non-binary position adjustment between an open valve position and a closed valve position.

[0100] The actuator configured to actuate the at least one valve in the generator heat management system may be configured to actuate the at least one valve in a manner that allows for position adjustment between an open valve position, one or more distinct intermediate valve positions, and a closed valve position. The non-binary position adjustment may therefore comprise a distinct position adjustment between at least three different valve positions. The actuator may be configured to actuate the valve between at least four defined positions. For example, the actuator may be configured to actuate the valve between four defined positions: open, closed, and two intermediate positions.

[0101] The actuator configured to actuate the at least one valve in the generator heat management system may be configured to actuate said valve so as to allow continuous position adjustment between an open valve position and a closed valve position.

[0102] 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.

[0103] The generator heat management system may be designed to control the temperature of a generator by cooling the oil used to cool and lubricate the generator, thereby transferring heat away from the generator.

[0104] The generator heat management system may include a plurality of heat exchangers; for example, one or more air-to-oil heat exchangers and one or more fuel-to-oil heat exchangers. The actuator may be configured to control fluid flow of a fluid through one or more of the heat exchangers, for example, an oil flow, an air flow, or a fuel flow.

[0105] The engine may include a plurality of generators. The generator heat management system may be configured to manage the temperature of the plurality of generators. The engine may include a plurality of generator heat management systems, each system being associated with a corresponding one of the plurality of generators.

[0106] The generator heat management system may include a plurality of valves. The at least one actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate at least two of the plurality of valves in the generator heat management system.

[0107] The generator heat management system may include a plurality of valves. The plurality of actuators may include a plurality of actuators configured to actuate at least one of the plurality of valves with the generator heat management system.

[0108] The engine may include an environmental bleed valve. The plurality of actuators may include an actuator configured to actuate the environmental bleed valve. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the actuator configured to actuate the environmental bleed valve.

[0109] The engine may include an engine heat management system. The plurality of actuators may include an actuator configured to actuate at least one valve in the engine heat management system. The fuel delivery system may be configured to provide fuel to hydraulically drive by fuel actuator configured to actuate the at least one valve in the engine heat management system.

[0110] The engine heat management system may include a plurality of valves.

[0111] The at least one actuator configured to actuate the at least one valve in the The engine heat management system may be configured to operate at least two of the plurality of valves in the engine heat management system.

[0112] The plurality of actuators includes a plurality of actuators each configured to actuate at least one of the plurality of valves in the engine heat management system.

[0113] The actuator configured to actuate the at least one valve in the engine heat management system may be configured to actuate the at least one valve in the engine heat management system in a manner to allow continuous position adjustment between an open valve position and a closed valve position.

[0114] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least ten, eleven, or at least twelve of the plurality of actuators.

[0115] The engine may include a plurality of actuated engine systems. The engine actuation systems may include a hydromechanical unit system and a turbine housing cooling system. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive at least three of the plurality of actuated engine systems.

[0116] The fuel supply system may be configured to supply fuel to hydraulically fuel-drive at least four, five, six, or seven of the powered engine systems.

[0117] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0118] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0119] The engine of the ninth aspect may be configured to perform the method of the second, fourth, sixth, and / or eighth aspects, and may have any of the characteristics described with respect to the first through eighth aspects.

[0120] According to a tenth aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where the bypass ratio, 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 at cruise conditions, is at least 4; a generator heat management system; a plurality of actuators, comprising an actuator configured to actuate a valve in the generator heat management system, wherein the actuator configured to actuate a valve in the generator heat management system is configured to actuate said valve so as to allow a non-binary position adjustment between an open valve position and a closed valve position; and a fuel supply system; and wherein the method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel drive at least the actuator of the generator heat management system valve of the plurality of actuators.

[0121] The position adjustment may be continuous, or may be distinct between three or more different positions, and possibly between four or more different positions,

[0122] The methods of the second, fourth, sixth, eighth, and tenth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the tenth aspect may be performed using the engine of the first aspect, the third aspect, the fifth aspect, the seventh aspect, or the ninth aspect.

[0123] It will be borne in mind that a feature described in relation to one aspect may be used in combination with any other aspect, mutatis mutandis.

[0124] As indicated elsewhere herein, the present disclosure may apply to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a bypass gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop engine, or a turbojet. Any such engine may or may not be provided with an afterburner. Such a turbine engine gas can be, for example, designed for land or marine power generation applications.

[0125] A gas turbine engine in accordance with any aspect of the present disclosure may include an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propeller).

[0126] Where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two counter-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propellers may rotate in opposite directions such that one rotates clockwise and the other counter-clockwise about the engine's rotational axis. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage arranged downstream of the propeller stage. The guide vane stage may be of variable pitch. Thus, the high-pressure, intermediate-pressure, and free-power turbines may respectively drive high- and intermediate-pressure propellers and compressors via suitable interconnecting shafts. Thus, the propellers may provide the majority of the propulsive thrust.

[0127] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reduction gear may be of the type described herein.

[0128] An engine according to the present disclosure may be a bypass engine. Such an engine may be a direct-drive bypass engine in which the fan is directly connected to the fan drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive bypass engine, the fan may be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0129] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan that is driven via a reduction gear. Accordingly, such a gas turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the reduction gear may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear. The core shaft may interlock the turbine and compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0130] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, e.g., one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

[0131] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (e.g., directly receive, e.g., via a generally annular conduit) a flow from the first compressor.

[0132] The reducer may be arranged to be driven by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above). For example, the reducer may be arranged to be driven only by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the reducer may be arranged to be driven by any one or more shafts, e.g. the first and / or second shafts in the example above.

[0133] The reducer may be a reduction gearbox (in that the output to the blower has a lower rotational speed than the input from the core shaft). Any type of reducer may be used. For example, the reducer may be a "planetary" or "star" reducer, as described in more detail elsewhere in this document. Such a reducer may be a single stage. Alternatively, such a reducer may be a compound reducer, for example a compound planetary reducer (which may have the input on the sun gear and the output on the ring gear, and hence be referred to as a "compound star" reducer), for example with two reduction stages.

[0134] The reducer may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2 or 3.2 to 3.8, for example in the range of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values ​​in the preceding sentence. Strictly by way of example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.1 or 3.2 to 3.8. Strictly by way of further example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.0 to 3.1. Strictly by way of example, the reducer may be a "planetary" reducer having a reduction ratio in the range of 3.6 to 4.2.In some arrangements the gear ratio may be outside these ranges.

[0135] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustor, which may be provided downstream of the fan and the compressor(s) with respect to the flow path (e.g., axially downstream). For example, the combustor may be directly downstream of (e.g., at the outlet of) the second compressor, when a second compressor is provided. As a further example, the flow at the outlet to the combustor may be supplied at the inlet of the second turbine, when a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0136] The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of attack may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct-drive bypass gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first (or "low pressure") compressor and either 10 or 11 stages in the second (or "high pressure") compressor. As a further example, the gas turbine engine may be a "geared" gas turbine engine (in which the blower is driven by a first core shaft via a reduction gearbox) comprising 11, 12 or 13 compressor stages (in addition to the blower). Such an engine may comprise 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As a further example, the gas turbine engine may be a "geared" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0137] The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from each other. The second (or "high pressure") turbine may comprise 2 stages in any arrangement (e.g., regardless of whether it is a geared or direct-drive engine). The gas turbine engine may be a direct-drive gas turbine engine comprising a first (or "low pressure") turbine having 5, 6, or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine comprising a first (or "low pressure") turbine having 3 or 4 stages.

[0138] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or a 0% span position, to a tip at a 100% span position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or in the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be included in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.32, or 0.29 to 0.30.These ratios can be commonly referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward-most) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e., the portion radially outboard of any platform.

[0139] The fan radius can be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The diameter of blower (which may simply be twice the blower radius) may be greater than (or of the order of) any of: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The blower diameter may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, within the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm.Strictly by way of non-limiting example, the blower diameter may be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm or 340 cm to 360 cm.

[0140] The rotational speed of the fan may vary during use. Generally, the rotational speed is lower for fans with a larger diameter. Strictly by way of non-limiting example, the rotational speed of the fan at cruise conditions may be less than 3500 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Strictly by way of further non-limiting example, the rotational speed of the fan at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 2,500 to 2,800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1,500 to 1,800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruising conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3,600 to 3,900 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

[0141] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a displacement of the fan blade tip with a velocity Utip. The work done by the fan blades on the flow results in an enthalpy increase dH of the flow. A fan tip load can be defined as dH / UpOint2, where dH is the enthalpy increase (e.g., the 1-D average enthalpy increase) across the fan and Utip is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity).The peak fan load at cruise conditions may be greater than (or in the order of) any one of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values ​​being dimensionless). The peak fan load may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.31 or 0.29 to 0.3 (e.g., for a geared gas turbine engine).

[0142] Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements the bypass ratio at cruise conditions may be greater than (or of the order of) any of the following: 9. 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 12 to 16, 13 to 15, or 13 to 14.Strictly by way of non-limiting example, 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. Strictly by way of further non-limiting example, 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 radially outboard of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0143] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before an inlet into the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or in the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 50 to 70.Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60.Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0144] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, specific thrust may depend, for a given thrust condition, on the specific fuel composition supplied to the combustion chamber. At cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any of the following: 110 Nkg 1 s, 105 Nkg 's, 100 Nkg 's, 95 Nkg 's, 90 Nkg 's, 85 Nkg 's or 80 Nkg 's. The specific thrust may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 80 Nkg 's to 100 Nkg 's, or of 85 Nkg 's to 95 Nkg 's. Such engines may be particularly efficient compared to conventional gas turbine engines.Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg 's to 95 Nkg 's. Strictly by way of non-limiting example, . the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg 's to 90 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg 's to 90 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 90 Nkg 's to 120 Nkg 's.

[0145] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN.Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN.The above mentioned thrust may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with engine static electricity.

[0146] In use, the temperature of the flow at the inlet of the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself may be referred to as blade nozzle director. In some examples, the TET may depend, for a given thrust condition, on the specific composition of fuel supplied to the combustion chamber. At cruise conditions, the TET may be at least (or in the order of) any of the following values: 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K, or 1650 K. Thus, solely by way of non-limiting example, the TET 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 1540 K to 1600 K. Strictly by way of non-limiting example, the TET 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 1590 K to 1650 K.Strictly by way of non-limiting example, the TET 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 1600K to 1660K. Strictly by way of non-limiting example, the TET 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 1590K to 1650K. Strictly by way of non-limiting example, the TET 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 1570K to 1630K.

[0147] The TET at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 1530 K to 1600 K. The maximum TET in engine operation may be, for example, at least (or in the order of) any one of the following: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, solely by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1960 K. Strictly by way of As a non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K.Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Strictly by way of non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET . may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range from 1800 K to 1950 K, or from 1900 K to 2000 K. The maximum TET may occur, for example, in a high thrust condition, for example, in a maximum takeoff thrust (MTP) condition.

[0148] A fan blade and / or airfoil portion of a fan blade described and / or claimed herein may be fabricated from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions fabricated using different materials.For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to withstand impact (e.g., by birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade may have a carbon fiber or aluminum-based (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0149] A fan as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be connected to the central portion in any desired manner. For example, each fan blade may comprise a fastening element that may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail that may notch into and / or engage a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.Any suitable method may be used to manufacture such a blade disc or blade ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the blades . blower can be connected to the hub / disc by welding, such as linear friction welding.

[0150] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the bypass duct outlet to be varied during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0151] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. When the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. When the fan blades have a metal body (for example, aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0152] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more parts thereof) have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, one skilled in the art would readily recognize that each term refers to all, or one or more parts, of a phase of engine operation within a given mission of an aircraft to which the gas turbine engine is designed to be attached.

[0153] As such, ground idling may refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a need for the engine to taxi. At idle, the engine may produce between 3% and 9% of the available engine thrust. In other non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. When taxiing, the engine may produce between 5% and 15% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In other non-limiting examples, the engine may produce between 7% and 10% of the available thrust.Takeoff may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In other non-limiting examples, the engine may produce 100% of the available thrust.

[0154] Climb may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In other non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to an operational phase of an aircraft's flight cycle between takeoff and arrival at cruise conditions, with arrival at cruise conditions thus defining the start of the cruise phase, or a portion thereof, of the aircraft's flight.Additionally or alternatively, climb may refer to a nominal point in, or one or more nominal periods during, an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may require an additional demand for engine thrust.

[0155] As used herein, cruise conditions, which may define the cruise phase (or a portion thereof) of aircraft flight, have a conventional meaning and are readily understood by those skilled in the art. In some examples, for a given gas turbine engine for an aircraft, cruise conditions may refer to the mid-cruise engine operating point of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise may be envisioned as the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the beginning of descent has been burned (which may be approximated by the midpoint—in terms of time and / or distance—between the end of climb and the beginning of descent).Cruise conditions may therefore define an operating point, a phase or part thereof, of the flight which provides thrust to ensure steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number), or at least substantially steady-state operation (i.e., maintaining at least a substantially constant altitude and / or at least a substantially constant Mach number) of an aircraft to which it is designed to be attached, taking into account the number of engines with which that aircraft is equipped. For example, where an engine is designed to be attached to an aircraft which has two engines of the same type, at cruise conditions the engine may provide half the total thrust which would be required for steady-state operation, or at least substantially steady-state operation, of that aircraft at mid-cruise.

[0156] In other words, for a given gas turbine engine for an aircraft, the cruise conditions can be defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady-state operation, or at least substantially steady-state operation, of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine at cruise conditions can be clearly defined.

[0157] Strictly by way of example, the forward speed at the cruise condition may be any point in the range of Mach 0.7 to 0.9, e.g., 0.75 to 0.85, e.g., 0.76 to 0.84, e.g., 0.77 to 0.83, e.g., 0.78 to 0.82, e.g., 0.79 to 0.81, e.g., of the order of Mach 0.8, of the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For a certain aircraft, the cruise conditions may be outside these ranges, e.g., below Mach 0.7 or above Mach 0.9.

[0158] Strictly by way of example, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

[0159] Strictly by way of example, cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of 20 kN to 40 kN.

[0160] Strictly by way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is of course dependent on the engine and its intended application and may be, for example, a value ranging from 35 kN to 65 kN.

[0161] In use, a gas turbine engine described and / or claimed herein may operate at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0162] Further, one skilled in the art would readily recognize that either or both of a descent and an approach refer to a phase of operation within an aircraft flight cycle between aircraft cruise and landing, with the approach in particular being part of the landing and takeoff (LTO) phase. During either or both of the descent and the approach, the engine may produce between 0% and 50% of the available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of the available thrust. In other non-limiting examples, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

[0163] According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect may correspond to an operating point, a phase, or a portion thereof, of the flight of the aircraft, as defined elsewhere herein.

[0164] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be at any suitable condition, which may be as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).

[0165] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may comprise (or may be) operation at any suitable condition, for example at mid-cruise of the aircraft, as defined elsewhere herein.

[0166] Those skilled in the art would understand that, unless mutually exclusive, any feature or parameter described in connection with any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter contained or described herein may be applied to any aspect and / or combined with any other feature or parameter described herein.

[0167] Furthermore, unless mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with any one or more other parameters and / or values ​​described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with any one or more other parameters and / or values ​​contained or described herein (e.g., one or more parameters B; one parameter C; and one parameter D, and so on) to express a product of their relationship. For example, those skilled in the art will understand that when parameter A is described separately from parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or any other application, combination, or function of parameter A with respect to parameter B, as appropriate.

[0168] Implementations will now be described by way of example only, with reference to the Figures, in which: [Fig.l] is a sectional side view of a gas turbine engine; [Fig.2] is a close-up sectional side view of an upstream portion of a geared gas turbine engine; [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine [Fig.4] is a close-up sectional side view of an upstream portion of a direct-drive gas turbine engine; [Fig.5] is a representation of an aircraft with a propulsion system comprising two gas turbine engines; [Fig.6] is a schematic representation of part of a fuel delivery system including a hydraulic fuel actuator; [Fig.7] is a schematic representation of part of a fuel delivery system comprising two hydraulic fuel actuators; [Fig.8] is a schematic representation of part of a fuel delivery system including a hydraulic fuel actuation system; [Fig.9] is a schematic representation of part of a fuel supply system comprising two hydraulic fuel actuation systems; [Fig. 10] is a schematic representation of an example of a fuel-driven hydraulic actuation system; [Fig. 11] is a schematic representation of another example of a fuel-operated hydraulic actuation system; [Fig. 12] is a schematic representation of another example of a fuel-operated hydraulic actuation system; [Fig. 13] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 14] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 15] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 16] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 17] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 18] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig. 19] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig.20] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig.21] is a schematic representation of another example of a fuel-driven hydraulic actuator; [Fig.22] is a schematic representation of an engine heat management system including an oil side valve operable using a fuel hydraulic actuator; [Fig.23] is a schematic representation of an engine heat management system including an air side valve operable using a fuel hydraulic actuator; [Fig.24] is a schematic representation of an engine heat management system comprising a plurality of heat exchangers; [Fig.25] is a schematic representation of another engine heat management system comprising a plurality of heat exchangers; [Fig.26] is a schematic representation of another reciprocating engine heat management system comprising an air-to-oil heat exchanger and a fuel-to-oil heat exchanger; [Fig.27] is a schematic representation of an example of a part of a fuel supply system with a valve operable using a hydraulic fuel actuator; [Fig.28] is a schematic representation of a generator heat management system including an oil side valve operable using a fuel hydraulic actuator; [Fig.29] is a schematic representation of a generator heat management system including an air side valve operable using a fuel hydraulic actuator; [Fig.30] is a schematic representation of a heat exchanger between engine oil and generator oil including actuable valves operable via fuel hydraulic actuators; [Fig.31] is a flowchart showing an example of a gas turbine engine operating process; [Fig.32] is a flowchart showing another example of a method of operating a gas turbine engine; [Fig.33] is a flowchart showing another example of a method of operating a gas turbine engine; [Fig.34] is a flowchart showing another example of a method of operating a gas turbine engine; and [Fig.35] is a flowchart showing another example of a gas turbine engine operating process.

[0169] The [Fig. 1] illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air intake 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The blower 23 is attached to, and driven by, the low pressure turbine 19 via a shaft 26 and an epicyclic reduction gear 30.

[0170] In use, the core air stream A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. Compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed to fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustion chamber 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive, the high-pressure and low-pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally acts to impart increased pressure to the bypass air stream B flowing through the bypass duct 22, such that the bypass air stream B is exhausted through the bypass exhaust nozzle 18 to generally provide the majority of the propulsive thrust. The epicyclic reduction gear 30 is a reduction box.

[0171] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in [Fig. 2]. The low pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outward of and meshing with the sun gear 28 are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation about the sun gear 28 in synchronism while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause it to rotate around the engine axis 9.Radially outwardly of and meshing with the planet gears 32 is a ring or crown gear 38 which is coupled, via links 40, to a stationary support structure 24.

[0172] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the reduction gear output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as an "intermediate pressure turbine" and "intermediate pressure compressor".When such alternative nomenclature is used, the blower 23 may be designated as the first compression stage or the lower pressure compression stage.

[0173] The epicyclic reduction gear 30 is shown by way of example in more detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic planetary reduction gear 30 generally include at least three planet gears 32.

[0174] The epicyclic reduction gear 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reduction gear 30 may be used. As a further example, the epicyclic reduction gear 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the reduction gear 30 may be a differential reduction gear in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0175] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the reducer 30 to the motor 10. As a further example, the connections (such as the links 36, 40 in the example of [Fig. 2]) between the reducer 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of rigidity or flexibility.As a further example, any suitable arrangement of bearings between rotating and stationary portions of the motor (e.g., between input and output shafts from the reducer and stationary structures, such as the reducer housing) may be used, and the description is not limited to the exemplary arrangement of [Fig. 2]. For example, where the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of the output and support links and bearing locations would typically be different from that shown as an example in [Fig. 2].

[0176] Thus, the present disclosure extends to a gas turbine engine having any arrangement of reduction gear styles (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing locations.

[0177] Optionally, the reducer may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster).

[0178] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have another number of compressors and / or turbines and / or another number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.l] has a split-flow nozzle 18, 20 which means that the flow through the bypass duct 22 has its own nozzle 18 which is independent of, and radially outward of, the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) may have a fixed or variable area.

[0179] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a reduction gear for the main shaft(s), being instead direct drive engines. A sectional view of such an engine is shown in [Fig.4].

[0180] With reference to [Fig.4], a gas turbine engine is generally indicated at 10, having a main axis of rotation 9. The engine 10 comprises, in axial flow series, an air intake 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0181] In use, air entering the intake 12 is accelerated by the blower 23 to produce two air streams: a core air stream A and a bypass air stream B. The core air stream A flows into the intermediate pressure compressor 14, and the bypass air stream B passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air stream A before delivering this air to the high pressure compressor 15 where further compression takes place.

[0182] Compressed air discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustion chamber 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive the high-pressure turbines, at pressure intermediate and low pressure 17, 19a, 19 before being discharged through nozzle 20 to provide additional propulsion thrust. The high pressure 17, intermediate pressure 19a and low pressure 19 turbines respectively drive the high pressure compressor 15, the intermediate pressure compressor 14 and the fan 23, each by a suitable interconnecting shaft.

[0183] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a reduction gear provided in the drive train from a turbine to a compressor and / or a fan.

[0184] While the described example relates to a bypass engine, the description may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[0185] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotation axis 9), a radial direction (in the bottom-up direction in [Fig.l]) and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial and circumferential directions are mutually perpendicular.

[0186] The fuel F supplied to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may comprise molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. Additionally or alternatively, when blended with, mixed with, or substituted for an alternative fuel, the fuel F may comprise renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F may comprise one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0187] By SAF, the skilled person means, for example, a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel or biojet fuel, produced from biological or non-biological resources. SAFs are commonly synthesized from carbon-containing gases extracted from the atmosphere and / or captured during industrial processes; or from a wide range of sustainable raw materials such as, for example, oils and used grease; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata, and moneygrass; non-biogenic alternative fuels; jatropha; halophytes and algae, rather than fossil-based hydrocarbons. SAF does not include fossil fuels.

[0188] The functional performance of a given fuel composition, or fuel mixture F for use in a given mission, may be defined, at least in part, by the fuel's ability to process the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and, emissions including gaseous and / or particulate matter. In this regard, particulate matter emissions may include soot particles created by the combustion of said fuel F, also known as non-volatile particulate matter (nvPM). Any reference to soot or smoke herein may also apply to other types of particulate matter emissions known in the art.The gaseous emissions may include one or more of nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO) including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and, volatile organic compounds (VOCs) created by the combustion of said fuel F. Any reference to gaseous emissions herein may also apply to other types of gaseous emissions known in the art.

[0189] A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, may at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, may at least partially reduce fuel takeoff volume, which may be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) may allow the fuel to experience elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can allow for reduced contrail formation, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance. For example, a relatively lower freezing point (°C) may allow for long-range missions. scope to optimize flight profiles; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to high aromatic fuels; and, maximum surface tension (mN / m) can ensure sufficient spray breakup and atomization of the fuel.

[0190] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels having higher ratios of hydrogen atoms to carbon atoms may have higher specific energies in the absence of bonding strain. For example, fossil-based hydrocarbon fuels may comprise molecules having about 7 to 18 carbons, with a significant portion of a given composition derived from molecules of 9 to 15 carbons, with an average of 12 carbons.

[0191] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with additional compositions pending approval. However, there is an anticipation in the aviation industry that sustainable aviation fuel blends, comprising up to (and including) 100% sustainable aviation fuel (SAF), will eventually be approved for use.

[0192] Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise lower aromatic and / or sulfur contents, relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise one and / or the other of higher iso-alkane content and cycloalkane content, relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may include a density between 90% and 98% of that of kerosene and / or a heating value between 101% and 105% of that of kerosene.

[0193] In some examples, the sustainable aviation fuel(s), or blend(s) supplied to the combustion equipment 16 may have a relatively lower aromatic and / or other non-paraffinic content than kerosene. Sustainable aviation fuel may have an aromatics content of, 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%. The aromatics content of the sustainable aviation fuel may be within an inclusive figure or range bounded by one or more of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005%, or 0%; or from 0% to 0.75%, from 0% to 0.5%, or from 0.1% to 0.25%; or from 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%; depending on one or more of a preference, a fuel stock or supplier, and compositional variations therein.

[0194] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide beneficial effects including, for example, one or more of higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx; and, lower CO2 emissions, compared to fossil-based hydrocarbon fuels (e.g., when combusted in combustion equipment 16).Thus, compared to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to either or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs.

[0195] As shown in [Fig. 5], an aircraft 1 may include a plurality of fuel tanks 50, 53a, b; for example, a larger primary fuel tank 50 located in the aircraft fuselage, and a smaller fuel tank 53a, 53b located in each wing. In other examples, an aircraft 1 may have only one fuel tank 50, and / or the wing fuel tanks 53a, b may be larger than the center fuel tank 50, or no center fuel tank 50 may be provided (all fuel being stored instead in the wings of the aircraft) - it will be appreciated that many different tank arrangements are contemplated and that the examples that have been illustrated are provided for ease of description and not intended to be limiting.

[0196] [Fig.5] shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The system The fuel supply system of the illustrated example comprises a single fuel source. As used herein, the term "fuel source" means either 1) a single fuel tank or 2) a plurality of fuel tanks that are fluidically interconnected. Each fuel source is configured to provide a separate source of fuel (i.e., a first fuel source may contain a first fuel having one or more characteristics different from a second fuel contained in a second fuel source). The first and second fuel sources are therefore not fluidically coupled to each other so as to separate the different fuels (at least under normal operating conditions).The use of multiple fuel sources allows an aircraft 1 to carry several different fuels, and to change the fuel used during operation, and possibly even during cruise or when changing from one operating stage to another during flight.

[0197] In the present examples, the first (and, in these examples, sole) fuel source comprises a center 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 port wing and at least one wing fuel tank is located in the starboard wing for balancing. All of the tanks 50, 53a, b are fluidically connected in the example shown, thereby forming a single fuel source. Each of the center fuel tank 50 and the wing fuel tanks 53a, b may comprise a plurality of fluidically interconnected fuel tanks.

[0198] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the center tank 50, thereby forming a second, separate fuel source. For balancing purposes, one or more port wing fuel tanks may be fluidly connected to one or more starboard wing fuel tanks. This may be done either through a center fuel tank (if that tank is not part of the other fuel source), or by bypassing the center fuel tank(s), or both (for maximum flexibility and safety). In another example, the first fuel source includes wing fuel tanks 53 and a center fuel tank 50, while a second fuel source includes another, separate center fuel tank.The fluidic interconnection between the wing fuel tanks and the center fuel tank of the first fuel source may be provided for balancing the aircraft 1. In aircraft 1 with multiple fuel sources, two or more of the fuel sources may therefore contain different fuels from each other, such that the aircraft 1 may change fuels during flight. It may therefore be more complex to . determining which fuel is supplied to the combustion chamber 16 rather than simply recording a single identity of a fuel on board the aircraft 1, or performing a check at start-up.

[0199] In some examples, the distribution of fuel tanks 50, 53 available on the aircraft 1 may be limited such that the first fuel source and the second fuel source are each substantially symmetrical about the centerline of the aircraft. In cases where an asymmetrical distribution of fuel tanks is permitted, a suitable means for transferring fuel is generally 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 center of mass of the aircraft can be maintained within acceptable lateral limits throughout the duration of the flight.

[0200] Aircraft typically refuel at several different airports, for example at the beginning and end of a long-haul flight. While there are standards that all aviation fuels must comply with, as discussed above, different aviation fuels have different compositions, for example depending on their source (e.g., different sources of petroleum, biofuels, or other synthetic aviation fuels (often described as sustainable aviation fuels, SAFs), and / or blends of petroleum-based fuels, and other fuels) and any included additives (e.g., such as antioxidants and metal deactivators, biocides, static reducers, icing inhibitors, corrosion inhibitors), and any impurities.The composition of available aviation fuel may not only vary from airport to airport and fuel supplier to fuel supplier, but also from batch to batch, even for a given airport or fuel supplier. Furthermore, the fuel tanks 50, 53 of the aircraft 1 are usually not emptied before being refilled for a subsequent flight, resulting in mixtures of different fuels in the tanks, with fuel of different composition actually being obtained from the mixture. One or more fuel ports 62 may be provided for refueling.

[0201] The fuel system 152 includes a main fuel flow path 110 from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10, and various branch fuel paths that take or return fuel to / from this main fuel flow path 110.

[0202] [Fig.6] is a schematic representation of a portion of a fuel delivery system 152 of a gas turbine engine 10. The main fuel flow path 110 of the fuel delivery system 152 serves as fuel source for fuel hydraulic systems. The gas turbine engine 10 includes a combustion chamber 16 and an actuator 154. The dotted arrows in [Fig.6] represent the direction of fuel flow.

[0203] The fuel supply system 152 is configured to supply fuel for combustion in the combustion chamber 16.

[0204] It will be appreciated that one or more fuel valves, pumps, sensors, taps for fuel hydraulic actuators, and heat exchangers may be present along the main fuel flow path 110, among other engine components. Fuel is supplied to the combustion chamber 16 directly via the line 110a, after passing through the most downstream engine component along the main fuel flow path 110.

[0205] Fuel delivered through lines 159 and 160 is used to drive (i.e., actuate) actuator 154. As such, actuator 154 is fuel hydraulically driven. The term "fuel hydraulic" as used herein refers to the hydraulic operation of an actuator whose hydraulic fluid is a fuel. Actuators with fuel hydraulic drive may be referred to herein as "fuel hydraulic actuators."

[0206] Fuel is delivered to the actuator 154 from the fuel flow path 110 via the line 159. After being used for the fuel hydraulic drive of the actuator 154, the fuel is returned to the main fuel flow path 110, via the line 160, and can then be delivered to the combustion chamber 16.

[0207] The fuel return line 160 may include a valve configured to regulate the return flow of fuel to the main fuel path 110, and the valve may be operated using a fuel hydraulic actuator. The fuel return line 160 may direct the fuel to a location along the main fuel flow path 110 that is before or after one or more heat exchangers of the main fuel flow path 110.

[0208] Once the fuel reaches the combustion chamber 16, it is burned to provide thrust, as explained in connection with [Fig.l].

[0209] [Fig. 6] shows a single fuel hydraulic actuator 154 on the fuel flow loop 159, 160. In other embodiments, multiple fuel hydraulic actuators 154 may be located on a single fuel flow loop 159, 160, utilizing fuel taken from a single point along the 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 the sample pipe 159 passes through all of the actuators, in turn) or in parallel with respect to the fuel flow (the sample pipe 159 passes through all of the actuators, in turn). 159 branching sample and one or more actuators located on each branch).

[0210] [Fig. 7] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a fuel source 50, 53, which may be a single fuel tank or multiple fuel tanks, and a fuel flow path 110 between the fuel source 50, 53 and the combustor 16. The illustrated fuel delivery system 152 includes two actuators 254a, b, each on a fuel evacuation loop separate from the main fuel flow path 110. The dotted arrows in [Fig. 7] represent the direction of fuel flow.

[0211] Both actuators 254a, b are hydraulically fuel driven. That is, both actuators are hydraulically fuel driven using fuel from source 50, 53 as fuel hydraulic fluid. Fuel passes from main fuel flow path 110 to actuators 254a, b respectively via lines 259a, b. Once fuel has been used to drive actuators 254a, b, it returns to the main fuel flow path via lines 260a, b.

[0212] Only two actuators 254a, b are shown in [Fig. 7] but, in other implementations, any plurality of actuators may be fuel hydraulically driven. In some embodiments, the gas turbine engine 10 may include a plurality of fuel hydraulically driven actuators and a plurality of actuators that are not fuel hydraulically driven.

[0213] [Fig. 7] shows a single fuel hydraulic actuator 254 on each fuel flow loop. In other embodiments, multiple fuel hydraulic actuators 254 may be located on one or more of the multiple fuel flow loops. The multiple actuators may be arranged in series with respect to the fuel flow (such that all of the fuel in the pickup pipe 159 passes through all of the actuators, in turn) or in parallel with respect to the fuel flow (with the pickup pipe 159 branching and one or more actuators located on each branch).

[0214] For some embodiments comprising a plurality of fuel hydraulic actuators, any fuel used for fuel hydraulic drive of the actuators may be taken from a single tapping point of the main fuel flow path (i.e., they may all be on a single fuel flow loop). In other embodiments, one or more of the plurality of fuel hydraulic actuators may be supplied with fuel from different tapping points of the main path. fuel flow (i.e. they may be on different fuel flow loops), but the fuel from each sampling point will be at the same pressure.

[0215] 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 may have its own fuel bleed pipe 159, 259 from the main fuel flow path 110, such that it has its own fuel flow loop. In other embodiments, two or more of these actuators may be located on the same fuel flow loop off the main fuel flow path 110.

[0216] 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 group of actuators used together to control a specific system, or perform a particular function, in the engine 10.

[0217] For example, the engine-powered motor systems of an engine 10 may include one or more of: • An engine heat management system (HMS); • A generator HMS; • A variable vane stator system (VSV); • A cabin environmental control system (e.g., a cabin air conditioning system); • A defrosting system; • A turbine case cooling (TCC) system; and / or • A landing chamber ventilation system.

[0218] Further, a given engine 10 may include multiple actuated systems of a given type, e.g., a plurality of deicing systems, each configured to deice a different component, multiple bearing chamber venting systems for different bearing chambers, and / or a backup cabin environmental control system.

[0219] [Fig. 8] is a schematic representation of a portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a main fuel flow path 110 from a fuel source 50, 53 to a combustion chamber 16, and an actuated system 356. The dotted arrows in [Fig. 8] represent the direction of fuel flow.

[0220] Similar to Figures 6 and 7, fuel is supplied to the combustion chamber 16, and fuel is supplied to the actuators 354b, 354c which are fuel hydraulic.

[0221] In the implementation of [Fig.8] the actuated system 356 comprises three actuators 354a-c. In other implementations, the actuated system 356 may comprise one actuator or any plurality of actuators.

[0222] In the implementation of [Fig.8], two actuators 354b, c are hydraulically driven by fuel. Each fuel hydraulic actuator 354b, c in this implementation has its own dedicated fuel flow loop. Multiple actuators may be supplied with fuel from the same tapping point of the main fuel flow path 110 in other implementations.

[0223] The remaining actuator 354a is not fuel hydraulically driven and may instead be actuated by any other suitable means (e.g., fuel hydraulically driven using oil, or electronically actuated, or pneumatically actuated. The dotted line indicates any means of actuation; whether the flow of a non-fuel liquid or electrical signals).

[0224] In some implementations, only one actuator of a given actuation system is fuel hydraulically driven. In other implementations, any plurality of actuators in a given actuation system are fuel hydraulically driven. All actuators in a given actuation system may be fuel hydraulically driven in some implementations. Any actuation system with at least one fuel hydraulic actuator may therefore be referred to as "fuel hydraulic."

[0225] [Fig. 9] is a schematic representation of a portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 is generally the same as that illustrated in [Fig. 8], but includes two actuated systems 456a, b. The dotted arrows in [Fig. 9] again represent the direction of fuel flow.

[0226] The first actuated system 456a includes three actuators 454a-c, two of which are fuel hydraulic. The second actuated system 456b includes two actuators 454d-e, both of which are fuel hydraulic. This implementation is given by way of example only and those skilled in the art will recognize that the turbine engine 10 may include any number of actuated systems, and that each actuated system may include any number of actuators. Any number of these actuators can be hydraulically driven by fuel.

[0227] The actuators described above may be any actuator used by the engine 10. Various specific actuators that may be hydraulically driven by fuel are described below. The actuation systems described herein may be any actuation system used by the aircraft. Various specific actuation systems that may be hydraulically driven by fuel are described below. It will be understood that the specific embodiments listed are described by way of example only, and the scope of protection is limited only by the claims.

[0228] An example of an actuated system is a variable stator vane (VSV) system. A VSV system controls the amount of airflow through the compressor 14, 15 (by moving one or more vanes; and typically by moving multiple stages of compressor vanes) to achieve optimal compressor performance. The VSV system varies the angle of the compressor stator vanes to manage compressor operability and efficiency.

[0229] In general, the variable stator vanes are designed to be rotatable about an axis to align with the angle of the incoming airflow. This movement of the vanes allows for a greater range of non-dimensional flow through the compressor 14, 15 without stall occurring; in effect, the compressor 14, 15 can operate smoothly over a greater range of engine power settings due to the VSVs. The VSVs are controlled by the engine control system in response to a range of parameters, generally including ambient conditions and the level of thrust demanded by the engine.

[0230] In engines 10 having multiple compressors 14, 15, a VSV system may be provided for one, more, or all of the compressors.

[0231] Many VSV systems include two VSV actuators, which can be used to change the angle of the compressor inlet guide vanes 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 implementations, one or more of the actuators of a VSV system are hydraulically driven by fuel.

[0232] [Fig. 10] shows a schematic representation of a fuel supply system 152 for an aircraft comprising a VSV system 556. The VSV system 565 comprises two actuators 554a, b. The fuel supply system 152 is configured to supply fuel to the combustion chamber 16 and to hydraulically fuel-drive the actuators 554a, b.

[0233] Another example of an actuated system is the turbine case cooling (TCC) system. Turbine engines 10 typically include a case surrounding the turbine 17, 19 and a TCC system is used to selectively cool this case. This cooling can be used to control rotor-stator tip clearances by reducing thermal expansion. A TCC system includes one or more valves operable to modulate the flow of bleed air from the engine's compressor 17, 19 across the case to provide cooling.

[0234] In engines 10 having multiple compressors 17, 19, a TCC system may be provided for one, some, or all of the turbines.

[0235] The valve(s) of the TCC system may be operated via an actuator. In various implementations, one or more of the actuators of a TCC system are hydraulically driven by fuel.

[0236] [Fig. 11] shows a schematic representation of a fuel supply system 152 for an aircraft 1 comprising a TCC system 656. The TCC system 656 comprises two actuators 654, 654', each actuator 654, 654' being configured to control a corresponding valve 654a, 654a'. The fuel supply system 152 is configured to supply fuel from the fuel source 50, 53 to the combustion chamber 16 and to hydraulically fuel-drive the actuators 654, 654'.

[0237] Another example of an actuated system used in the turbine engines 10 is a hydromechanical unit (HMU). The HMU primarily distributes and regulates the flow of fuel to the combustion chamber 16. It may also supply fuel to the actuators.

[0238] The HMU includes a plurality of actuators configured to operate a plurality of valves. In various implementations, one or more of these actuators are hydraulically driven by fuel.

[0239] Hydromechanical units typically comprise 3 or 4 actuators. In various implementations, one, some, or all of these actuators may be hydraulically driven by fuel.

[0240] The hydromechanical unit may include an actuator configured to regulate the flow of fuel to the combustion chamber 16 (e.g., it may be based on parameters including engine thrust demand and fuel heating value), which includes controlling the spillage of fuel around one or more fuel pumps of the engine (not all of the fuel exiting a pump is necessarily supplied to the combustion chamber 16; some may be recirculated ("spilled"), and the recirculated fuel generally constitutes a proportion of the fuel entering the pump inlet. The spilled fuel can therefore be used to perform engine functions as well as to allow a pump to continue operating at a set rate even in the event of fluctuations in combustion chamber fuel demand. As used herein, "spilled" fuel therefore includes fuel used for any purpose other than supplying the combustion chamber 16, e.g., for the fuel hydraulic circuit, not just fuel sent directly from the pump outlet to the pump inlet).

[0241] The hydromechanical unit may include an actuator configured to shut off the fuel supply. The hydromechanical unit may include an actuator configured to provide minimum fuel pressure or control a fuel pressure drop. The hydromechanical unit may include an actuator configured to control the fuel supply through multiple pumps. The hydromechanical unit may include an actuator configured to open a path to a drain manifold when the engine is shut down (e.g., for fuel return procedures that may use a small secondary fuel tank installed on the engine 10 rather than the main tank(s) 50, 53 of the aircraft). The hydromechanical unit may include an actuator associated with VSV control. The hydromechanical unit may include an actuator associated with a thrust control housing.The hydromechanical unit may include an actuator configured to split fuel delivery (e.g., to distribute fuel to different manifolds to supply injectors to the combustion chamber 16, and / or for different hydraulic fuel controls). In some implementations, TCC control and / or variable bleed valve control may be performed using one or more actuators located in the HMU (so there may be some overlap of the actuated systems).

[0242] In some implementations, the engine 10 may include a hydromechanical metering unit, which may be located along the primary fuel flow path 110 through the engine, upstream of a fuel metering valve that controls the flow rate of fuel into the combustion chamber 16. In such implementations, the, or each, fuel sampling point may be located adjacent to or within the hydromechanical metering unit.

[0243] [Fig. 12] shows a schematic representation of a fuel supply system 152 for an aircraft 1 comprising an HMU 756. The HMU 756 shown comprises three actuators 754, 754', 754”. The fuel supply system 152 is configured to supply fuel from the fuel source 50, 53 to the combustion chamber 16 and to hydraulically fuel-drive the actuators 754. Each actuator 754 is configured to control an associated valve 754a, 754a', 754a”. [Fig. 12] shows each actuator 754 with its own dedicated sampling point on the main fuel flow path 110, and its own dedicated fuel loop. In other implementations, a branch path and / or a distribution manifold may be provided to split fuel from a single sampling point on the main fuel flow path 110 between multiple hydraulic fuel actuators. A hydromechanical measurement unit, if desired, may be integrated into the HMU.

[0244] Another example of an actuator used in turbine engines 10 is a vent valve actuator (also referred to herein as "vent valve actuator"). Vent valve actuators control the opening and closing of vent valves, which control the opening and closing of vent lines for one or more engine bearing chambers.

[0245] Various implementations may include one or more ventilation valves controlled by one or more ventilation valve actuators. In various implementations, one or more of these ventilation valve actuators are hydraulically driven by fuel.

[0246] [Fig. 13] shows a schematic representation of a fuel supply system 152 for an aircraft 1 including a vent valve 854a. The vent valve 854a is controlled via an actuator 854. The fuel supply system 152 is configured to supply fuel from a fuel source 50, 53 to hydraulically fuel-drive the actuator 854 and supply fuel to the combustion chamber 16. The actuated opening and closing of the vent valve 854a controls the flow of air between the engine bearing chamber and the atmosphere. A vent valve system may be defined, including at least the vent valve 854a and its actuator 854, and optionally including multiple valves and / or actuators.

[0247] Another example of an actuated system used in the turbine engines 10 is a bleed air system. The bleed air systems 954, 954a, 1054, 1054a, 1154, 1154a, each including at least one valve and its associated actuator, redirect bleed air (compressed air taken from the compressor 14, 15 upstream of the combustion chamber 16) for use in other systems. The bleed air is useful in various systems due to its relatively high temperature and pressure. The bleed air systems may be referred to as "engine bleed air systems" (EBAS).

[0248] Many air bleed systems may be used in a turbine engine 10, some of which are described below. The air bleed systems include one or more valves controlled by one or more actuators. In various implementations, one or more of the actuators of an air bleed system are hydraulically driven by fuel.

[0249] The air bleed systems used to control the pressure in the cabin of the aircraft 1 are called environmental air bleed systems.

[0250] Air bleed systems may be used to eliminate or prevent ice formation on the nacelle 21 by redirecting high pressure, high temperature air from the compressor 14, 15 onto the nacelle 21. Such systems include a nacelle anti-ice valve 954a that is controlled by an actuator 954. In various implementations, this actuator is hydraulically driven by fuel.

[0251] [Fig. 14] shows a schematic representation of a fuel supply system 152 including a fuel source 50, 53 (grouped into fuel flow path 110 for clarity) and a combustor 16. The fuel supply system 152 is configured to supply fuel to the combustor 16 and to supply the fuel to an actuator 954 for hydraulically fuel-driving the actuator 954. The actuator 954 is configured to actuate a valve 954a that allows passage of bleed air onto the nacelle 21 of a gas turbine engine 10.

[0252] The air bleed systems may also be used to remove or prevent ice formation on the engine stators by redirecting high pressure, high temperature air from the compressor 14, 15 onto the engine stators. Such systems may include an engine section stator anti-ice (ESS) valve that is controlled by an actuator. In various implementations, this actuator is hydraulically driven by fuel.

[0253] [Fig. 15] shows a schematic representation of a fuel supply system 152 comprising a fuel source 50, 53 and a combustion chamber 16. The fuel supply system 152 is configured to supply fuel to the combustion chamber 16 and to supply the fuel to a bleed air actuator 1054 for hydraulically fuel-driving the actuator 1054. The actuator 1054 is configured to operate a valve 1054a that allows passage of bleed air over / over one or more stators of a turbine engine 10 (the valve 1054a may be an ESS anti-icing valve).

[0254] Bleed air systems may also be used to control the air conditioning in the cabin of an aircraft 1. In these systems, the bleed air is filtered and cooled (by means of expansion and / or heat exchange, for example, with atmospheric air) using a conditioning unit (conditioning units or pressurization air conditioning kits (PACKS) are two accepted terms). Air conditioning units are used to cool and, if necessary, dehumidify the bleed air from the engine or of the APU before it is sent into the aircraft cabin. Such systems include one or more valves and one or more actuators. Air from the engine 10 is generally routed to the unit through a one-way check valve 1154a. In various implementations, one or more of these actuators are hydraulically driven by fuel.

[0255] [Fig. 16] shows a schematic representation of a fuel supply system 152 configured to supply fuel to the combustion chamber 16 and to supply fuel to a conditioning unit actuator 1154 to hydraulically fuel-drive the actuator 1154. The actuator 1154 is configured to operate a valve 1154a that allows the passage of bleed air into a conditioning unit. The valve 1154a may therefore be referred to as an environmental air bleed valve; an air conditioning unit valve is one example.

[0256] Extraction of bleed air from the compressor for air conditioning, prior to use in other systems, is typically accomplished via an upper stage valve (also known as a "high pressure valve") that is controlled by an actuator. In various implementations, this actuator is hydraulically driven by fuel. The bleed air extracted from the compressor 14, 15 via the upper stage valve is then redirected via one or more additional valves for use in aircraft cabin environmental control.

[0257] [Fig. 17] shows a schematic representation of a fuel supply system 152 configured to supply fuel to the combustion chamber 16 and to an upper stage valve actuator 1254 to hydraulically fuel-drive the actuator 1254. The actuator 1254 is configured to operate an upper stage valve 1254a which allows extraction of bleed air from the compressor 14, 15 for subsequent use.

[0258] Bleed air may be extracted from other parts of the engine 10 and / or other parts of the aircraft 1, in addition to, or instead of, the compressor 14, 15 (e.g., bleed air may be taken from a bypass duct 22 or from an APU). Extraction of this bleed air is via an engine bleed air valve (also called a "manifold pressure valve" or pressure control valve) which is controlled by an actuator. In various implementations, this actuator is hydraulically driven by fuel.

[0259] The bleed air can typically come from three sources: an intermediate pressure bleed (IP compressor 14), a high pressure bleed (HP compressor 15), or a bypass duct 22 (i.e., downstream of the blower 23).

[0260] Each bleed air source may be associated with a valve. In addition, there is typically a valve that regulates the static pressure (downstream of the IP and HP sources) according to the requirements of the end user of the aircraft systems. Each of these valves may be hydraulically actuated by fuel.

[0261] The manifold pressure valve is used to regulate the pressure of the bleed air according to its desired use(s).

[0262] [Fig. 18] shows a schematic representation of a fuel supply system 152 configured to supply fuel to the combustion chamber 16 and to an actuator 1354 for hydraulically fuel-driving the actuator 1354. The actuator 1354 is configured to operate a valve 1354a that allows extraction of bleed air from parts of the engine other than the compressor, for example, from a bypass duct 22.

[0263] Bleed air may be extracted from an auxiliary power unit (APU) 1455; the APU may be located in the engine 10 in some implementations, or elsewhere in the aircraft in other implementations. Extraction of bleed air from the APU is typically via an APU bleed air valve that is controlled by an actuator. In various implementations, this actuator is hydraulically driven by fuel. The dotted arrow in [Fig. 19] depicts a flow of bleed air from the APU 1454' to an APU bleed air valve 1454a.

[0264] In some implementations, the fuel used to drive the APU bleed valve may come from a supply line located between the fuel source and the engine, while the fuel used for other fuel hydraulic actuators may come from supply lines in the engine. The APU bleed valve 1454a, and even the APU 1454', may therefore not be part of the engine 10 in some implementations.

[0265] [Fig. 19] shows a schematic representation of a fuel delivery system 152 configured to supply fuel to the combustion chamber 16 and to an APU bleed air actuator 1454 to hydraulically fuel-drive the actuator 1454. The actuator 1454 is configured to operate a valve 1454a that allows extraction of bleed air from an auxiliary power unit.

[0266] Bleed air from the APU 1455 may be used to facilitate various tasks, such as starting the engine 10. An engine start valve controls the flow of bleed air from the APU 1455 to the turbine 17, 19. The engine start valve is controlled by an actuator. In various implementations, this actuator is hydraulically driven by fuel.

[0267] [Fig.20] shows a schematic representation of a fuel delivery system 152 configured to supply fuel to the combustion chamber 16 and to an engine start valve actuator 1554 to hydraulically fuel-drive the actuator 1554. The actuator 1554 is configured to operate an engine starter valve 1554a that redirects bleed air from the auxiliary power unit for use in starting the turbine engine 10.

[0268] Bleed air from one turbine engine 10 may be introduced into components of another turbine engine, for an aircraft with multiple engines 10 (e.g., on the opposite wing of an aircraft 1). Bleed air flow between engines 10 is regulated by an isolation valve 1654a. The isolation valve is controlled by an actuator 1654. In various implementations, this actuator is hydraulically driven by fuel.

[0269] [Fig.21] shows a schematic representation of a fuel delivery system 152 configured to supply fuel to the combustion chamber 16 and to an isolation valve actuator 1654 for hydraulically fuel-driving the actuator 1654. The actuator 1654 is configured to operate an isolation valve 1654a.

[0270] Another example of an actuated system used in turbine engines 10 is a heat management system (HMS). Turbine engines 10 often include an engine HMS and a generator HMS. Both HMS systems control the oil temperature by enabling and controlling the use (or bypass) of the air-to-oil (and / or fuel-to-oil) heat exchangers.

[0271] The engine HMS system controls the temperature of the oil system that lubricates the engine components (e.g., bearing chambers, reduction gear, etc.). The engine HMS system includes one or more engine HMS valves that are controlled by one or more actuators. In various implementations, one or more of these actuators are hydraulically driven by fuel.

[0272] For example, in some implementations, the engine HMS includes a single valve and a single fuel-driven hydraulically driven actuator for controlling the valve.

[0273] The engine HMS system may include multiple heat exchangers arranged in a parallel or series configuration. A single valve controlled by a single fuel hydraulic actuator may allow the use (or bypass) of all of the heat exchangers (in particular, the flow rate of a fluid (air, oil, or fuel, if applicable) may be controlled by the valve, with the proportion of fluid not passing through the heat exchanger being directed along a bypass pipe. In some implementations, each heat exchanger is associated with a valve controlled by an associated actuator. In various implementations, one or more of these actuators are hydraulically driven by fuel. In addition, one or more recirculation pipes may be provided to allow fluid from an outlet of the heat exchanger to be returned to the inlet of the heat exchanger, thereby passing through the heat exchanger several times. The flow rate in the recirculation pipe(s) may again be adjusted using a valve controlled by an actuator. In some implementations, a pump, or other fuel delivery means, may be in, or associated with, the recirculation pipe(s) to drive this recirculation.

[0274] [Fig.22] shows a schematic view of a section of an engine HMS 1700. The engine HMS includes an oil-to-air heat exchanger 1703 configured to remove heat from the engine oil. The engine HMS 1700 includes an oil line 1701 and an air line 1702.

[0275] The engine HMS 1700 includes a valve 1704a positioned in the oil line 1701. The valve 1704a is configured to direct oil through the heat exchanger 1703 or to direct oil to bypass the heat exchanger 1703. The valve 1704a is controlled via an actuator 1704 that is hydraulically driven by fuel.

[0276] The actuator 1704 is configured to actuate the valve 1704a so as to allow non-binary position adjustment between an open valve position and a closed valve position (i.e., the actuator 1704 may allow analog rather than binary control of the valve 1704a, or it may allow the valve to be adjusted to one or more intermediate positions between open and closed). The valve 1704a and the associated fuel hydraulic actuator 1704 are thus configured to control the amount of oil passing through the heat exchanger 1703 to regulate the cooling of the oil.

[0277] [Fig. 23] shows a schematic view of a section of an engine HMS 1700 according to an alternative implementation. The engine HMS 1700 includes an oil-to-air heat exchanger 1703 configured to remove heat from the engine oil. The engine HMS 1700 includes an oil line 1701 and an air line 1702, as for the implementation described with respect to [Fig. 22]. The fuel hydraulically controlled valve 1704a' is, however, located differently.

[0278] The engine HMS 1700 of [Fig.23] includes a valve 1704a' positioned in the air line 1702. The valve 1704a' is configured to direct air through the heat exchanger 1703 or to direct air to bypass the heat exchanger 1703. The valve 1704a' is controlled via an actuator 1704' that is hydraulically driven by fuel.

[0279] The actuator 1704' is again configured to actuate the valve 1704a' so as to allow non-binary position adjustment between an open valve position and a closed valve position (i.e., the actuator 1704' may allow analog rather than binary control of the valve 1704a', or it may allow the valve to be adjusted to one or more intermediate positions between open and closed). The valve 1704a' and the associated fuel hydraulic actuator 1704' are thus configured to control the amount of air passing through the heat exchanger 1803 to regulate the cooling of the oil.

[0280] In other implementations, the engine HMS of Figures 22 and 23 may be combined such that the HMS includes a valve 1704a, 1704a' actuated via the fuel hydraulic actuators 1704, 1704' in both the air line and the oil line.

[0281] In Figures 22 and 23, the heat exchanger 1703 is a parallel flow heat exchanger. In other implementations, the heat exchanger 1703 may be a counterflow heat exchanger, meaning that the direction of flow of the air line or the oil line is reversed.

[0282] [Fig. 24] shows another implementation in which the engine HMS 1700 includes three heat exchangers 1703a-c arranged in parallel with respect to the oil flow (located on branches of the oil line 1701). Although the implementation of [Fig. 24] includes three heat exchangers, it is possible to use any plurality of heat exchangers in other implementations. The three heat exchangers 1703a-c are all air-oil heat exchangers in the illustrated implementation.

[0283] The engine HMS 1700 includes an oil line 1701 that supplies oil to the heat exchangers 1703a-c. The engine HMS 1700 includes three air lines 1702a-c that respectively supply air to the heat exchangers 1703a-c.

[0284] The engine HMS 1700 includes a valve 1704a in the oil line 1701 that is configured to allow oil to pass through, or bypass, all of the heat exchangers 1703a-c. The valve 1704a is controlled via a fuel-driven hydraulic actuator 1704.

[0285] The actuator 1704 is configured to actuate the valve 1704a so as to allow non-binary position adjustment between an open valve position and a closed valve position (i.e., the valve may be continuously adjustable, or adjustable between three or more positions). The valve 1704a and the associated fuel hydraulic actuator 1704 are thus configured to control the amount of oil passing through the heat exchangers 1703a-c to regulate the cooling of the oil.

[0286] In other implementations, the engine HMS 1700 includes additional valves, each valve being disposed in an air line 1702a-c and being configured to allow air to bypass the corresponding air-to-oil heat exchanger 1703a-c. One or more of the valves disposed in the air lines 1702a-c may be hydraulically driven by fuel.

[0287] [Fig. 25] shows another implementation in which the engine HMS 1700 again includes three heat exchangers 1703a-c arranged in parallel, but in which the valve arrangement is different from that of [Fig. 24]. Although the implementation of [Fig. 25] includes three heat exchangers, it is possible to use any plurality of heat exchangers in other implementations.

[0288] The engine HMS 1700 includes an oil line 1701 that supplies oil to the heat exchangers 1703a to c. The engine HMS 1700 includes three air lines 1702a to c that respectively supply air to the heat exchangers 1703a to c

[0289] The engine HMS 1700 includes three valves 1704a, 1704a', 1704a” (collectively, 1704a), each valve 1704a being positioned in the oil line 1701, each on a different branch thereof. Each valve 1704a is configured to allow oil to pass through, or bypass, an associated heat exchanger 1703a-c. The valves 1704a allow oil to bypass one or more of the heat exchangers 1703a-c to control the rate and amount of heat removal 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 may be hydraulically driven by fuel.

[0290] The actuators 1704 are configured to actuate the valves 1704a so as to allow non-binary position adjustment between an open valve position and a closed valve position for each valve 1704a (i.e., the valve may be continuously adjustable, or adjustable between three or more positions). The valves 1704a and the associated fuel hydraulic actuators 1704 are thus configured to individually control the amount of oil passing through each of the heat exchangers 1703a-c to regulate the cooling of the oil.

[0291] In other implementations, the engine HMS 1700 includes additional valves, each valve being disposed in an air line 1702a-c and being configured to allow air to bypass the corresponding air-to-oil heat exchanger 1703a-c. One or more of the valves disposed in the air lines 1702a-c may be hydraulically driven by fuel.

[0292] In other implementations, the engine HMS systems 1700 of Figures 24 and 25 may be combined such that the HMS system includes a valve allowing simultaneous bypass of all heat exchangers as well as valves for bypassing individual heat exchangers.

[0293] [Fig.26] shows a schematic view of an engine HMS 1700 including a fuel-oil heat exchanger 1703a and an air-oil heat exchanger 1703b. The engine HMS 1700 includes an oil line 1701 including a valve 1704a. The valve 1704a is operable via a fuel-driven hydraulic actuator 1704. The valve 1704a divides the oil line 1701 into three portions, or branches, and regulates the flow of oil in each portion.

[0294] The first portion 1701a of the oil line passes through the fuel-oil heat exchanger 1703a. The fuel passes through the fuel-oil heat exchanger 1703a via the fuel line 1705. The fuel in the fuel line 1705 that passes through the fuel-oil heat exchanger 1703a may be used to fuel-hydraulic drive the actuator 1704 to control the valve 1704a. The fuel may be used to fuel-hydraulic drive the actuator 1704 to control the valve 1704a after the fuel passes through the heat exchanger 1703a, or before the fuel passes through the heat exchanger 1703a.

[0295] The second part of the oil line 1701b does not pass through any of the heat exchangers 1703a,b.

[0296] The third part of the oil line 1701c passes through the air-oil heat exchanger 1703b. The air passes through the air-oil heat exchanger 1703b via the air line 1702.

[0297] The engine HMS 1700 of [Fig. 26] may include one or more additional valves in the oil line 1701c and / or the oil line 1701a. These valves may be hydraulically driven by fuel. The valves may allow oil to bypass the associated heat exchangers. In such implementations, the oil line 1701b may be omitted. In other implementations, one or more additional valves may be included in the air line 1702 and / or the fuel line 1705 to allow air or fuel (if applicable) to bypass the respective heat exchanger 1703a,b.

[0298] [Fig.27] shows a schematic view of a section of a fuel delivery system 152. The fuel delivery system 152 includes a valve 1804a configured to split the fuel into two paths. The valve 1804a may be actuated using a hydraulically fuel-driven actuator 1804. The first path 1801 diverts the fuel upstream of a fuel-oil heat exchanger 1803 such that the fuel passes through the heat exchanger fuel-oil flow path 1803 as it travels to the combustion chamber 16. The second path 1802 diverts the fuel downstream of the fuel-oil heat exchanger 1803 so that the fuel is not heated as it travels to the combustion chamber 16. This section of the fuel system 152 can be used to return the fuel to the main fuel flow path 110. The valve 1804a can be used to decide where along the fuel path to divert how much fuel. The fuel-oil heat exchanger 1803 can be the fuel-oil heat exchanger 1703a of the implementation of [Fig. 26], or a different fuel-oil heat exchanger.

[0299] It will be appreciated that fuel from the one or more fuel flow loops 159, 259, etc., for hydraulic fuel actuation may be returned to the main fuel flow path 110 at any convenient location (in some cases, the fuel may even be returned to the fuel tank 50, 53, although it is more common for the fuel to be retained in the engine 10 and simply recirculated within the engine 10 until it is sent to the combustion chamber 16 for combustion).The ability to control where fuel is returned along the primary fuel flow path 110 may contribute to overall engine heat management: for example, if the temperature of the fuel leaving the one or more actuators is below a threshold, it may be returned to a point along the flow path 110 before a fuel-to-oil heat exchanger, to allow the fuel temperature to be further increased before it reaches the combustion chamber 16. In contrast, if the temperature of the fuel leaving the one or more actuators is above a threshold, it may be returned to a point along the flow path 110 after a fuel-to-oil heat exchanger, since no further heat input to the fuel is desired, and possibly also to a point downstream of an engine pump to reduce the risk of thermal damage to the pump.

[0300] In implementations with multiple heat exchangers along the main fuel flow path 110, the valve 1804a may be configured to direct fuel through one or more of the heat exchangers or bypass one or more of them.

[0301] The generator HMS 1900 may be independent of the engine HMS 1700. It provides cooling for an electrical machine used, for example, to power the airframe with electricity to operate the aircraft systems. The generator HMS 1900 includes one or more generator HMS valves 1904a, 1904a' that are controlled by one or more actuators 1904, 1904'. In various implementations, one or more of these actuators 1904, 1904' are hydraulically driven by fuel.

[0302] For example, in some implementations, the generator HMS 1900 includes a single valve and a single actuator that is hydraulically driven by fuel.

[0303] The generator HMS system 1900 may include a plurality of heat exchangers arranged in a parallel configuration. A single valve controlled by a single fuel hydraulic actuator may allow use (or bypass) of the heat exchangers. In some implementations, each heat exchanger is associated with a valve controlled by an associated actuator. In various implementations, one or more of these actuators are fuel hydraulically driven.

[0304] [Fig.28] shows a schematic view of a section of a generator HMS 1900. The generator HMS 1900 includes an oil-to-air heat exchanger 1903 configured to remove heat from the engine oil, and more specifically to remove heat from the oil used to cool (and possibly also to lubricate) the generator. The generator HMS 1900 includes an oil line 1901 and an air line 1902.

[0305] The generator HMS 1900 includes a valve 1904a positioned in the oil line 1901. The valve 1904a is configured to direct oil through the heat exchanger 1903 or to direct oil to bypass the heat exchanger 1903. The valve 1904a is controlled via an actuator 1904 that is hydraulically driven by fuel.

[0306] The actuator 1904 is configured to actuate the valve 1904a so as to allow continuous position adjustment between an open valve position and a closed valve position, thereby continuously adjusting the flow rate of oil through the air-oil heat exchanger 1903. The valve 1904a and the associated fuel hydraulic actuator 1904 are therefore configured to control the amount of oil passing through the heat exchanger 1903 to regulate the cooling of the oil.

[0307] In other implementations, the oil-to-air heat exchanger 1902 may be replaced by, or used in addition to, a fuel-to-oil heat exchanger. In such implementations, the actuator 1904 may be hydraulically fuel-driven by fuel before said fuel is heated by oil in the heat exchanger. Alternatively, the actuator 1904 may be hydraulically fuel-driven by fuel after said fuel has been heated by oil in the heat exchanger.

[0308] A valve and associated fuel-hydraulic controllable actuator may be provided on the air line (for an oil-to-air heat exchanger) or on the fuel line (for an oil-to-fuel heat exchanger), in addition to, or instead of, the valve provided on the oil line, in various implementations.

[0309] [Fig.29] shows a schematic view of a section of a generator HMS 1900 according to such an alternative implementation. The generator HMS 1900 includes an oil-to-air heat exchanger 1903 configured to remove heat from the engine oil and more specifically to remove heat from the oil used to cool (and possibly also to lubricate) the generator. The generator HMS 1900 includes an oil line 1901 and an air line 1902.

[0310] The generator HMS 1900 includes a valve 1904a' positioned in the air line 1902. The valve 1904a' is configured to direct air through the heat exchanger 1903 or to direct air to bypass the heat exchanger 1903. The valve 1904a' is controlled via an actuator 1904' that is hydraulically driven by fuel.

[0311] The actuator 1904' is configured to actuate the valve 1904a' so as to allow continuous position adjustment between an open valve position and a closed valve position. The valve 1904a' and the associated fuel hydraulic actuator 1904' are thus configured to control the amount of air passing through the heat exchanger 1903 to regulate the cooling of the oil.

[0312] In other implementations, the generator HMS of Figures 28 and 29 may be combined such that the generator HMS 1900 includes a valve actuated via a fuel hydraulic actuator in both the air line and the oil line.

[0313] In Figures 28 and 29, the heat exchanger 1903 is a parallel flow heat exchanger. In other implementations, the heat exchanger 1903 may be a counterflow heat exchanger, meaning that the direction of flow of the air line or the oil line is reversed.

[0314] [Fig.30] shows a schematic view of a section of a combined HMS 2000. The combined HMS 2000 includes an oil-to-oil heat exchanger 2005 configured to allow heat transfer between the oil in the generator HMS 1900 (which may be referred to as generator oil, for brevity) and the oil in the engine HMS 1700 (which may be referred to as main engine oil, for brevity). Heat can therefore be transferred between the two systems, effectively combining them into a single combined heat management system, while keeping the oils fluidly isolated.

[0315] In [Fig.30] the heat exchanger 2005 is a parallel flow heat exchanger. In other cases, the direction of flow of one or the other fluid in the heat exchanger may be reversed and the heat exchanger 2005 may therefore be a counterflow heat exchanger.

[0316] The main engine oil enters the heat exchanger 2005 via the oil line 1701 of the engine HMS 1700. In some implementations, the oil line 1701 includes a valve 1704 configured to divert some or all of the oil to bypass the heat exchanger 2005. This valve 1704 may be actuated using a fuel-driven hydraulic actuator 1704'.

[0317] Generator oil enters heat exchanger 2005 via oil line 1901 of generator HMS 1900. In some implementations, oil line 1901 includes a valve 1904a” configured to divert some or all of the oil to bypass heat exchanger 2005. This valve 1904a” may be actuated using an actuator 1904” that is hydraulically driven by fuel.

[0318] Various methods of operating a gas turbine engine for an aircraft are described below.

[0319] [Fig. 31] is a flowchart showing a method 1000 of operating a gas turbine engine 10 according to various implementations. The method 1000 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and 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 actuators 254; and a fuel system 152.

[0320] In the described implementations, a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass conduit 22 to the mass flow rate of the flow through the core 11 under cruising conditions, is greater than or equal to 4. Such bypass ratios may be used in any of the engines used to perform any of the methods of FIGS. 31-35.

[0321] The method 1000 comprises: supplying 1010, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and providing 1020, using the fuel supply system 152, fuel to hydraulically fuel-drive at least ten of the plurality of actuators 254.

[0322] The at least ten actuators 254 may comprise: • One or other of the at least two valves 554a, b of the VSV system 556; • Either of the two valves 654a, 654a' of the TCC 656 system; • Any one of the at least three valves 754a, 754a', 754a” of the HMU 756; • The ventilation valve actuator 854; • The 954 nacelle anti-icing valve actuator or any other anti-icing valve actuator; • The air sampling actuator 1054; • The conditioning group actuator 1154; • The upper stage valve actuator 1254; • Actuator 1354 for controlling the bleed air extraction valve from parts of the engine that are not compressors; • The APU 1454 air bleed actuator; • The 1554 engine start valve actuator; • The 1654 isolation valve actuator; • Any of the actuators 1704, 1704' of the engine heat management system 1700; and / or • Any of the actuators 1904, 1904' of the generator heat management system 1900.

[0323] The fuel used to hydraulically fuel one or more of the actuators 254 may then be delivered to the combustion chamber 16 for combustion. The fuel may flow along a primary fuel flow path 110 to the combustion chamber 16; the primary fuel flow path 110 may then deliver 1010 the fuel to the combustion chamber 16. A portion of the fuel flowing along the primary fuel flow path 110 may be diverted from the primary fuel flow path 110 to the actuators 254; this fuel may then be returned to the primary fuel flow path 110 after being used for actuation. One or more hoses 159, 160 may be used to take fuel from the main fuel flow path 110 to the actuators 254. These hoses may form one or more fuel flow loops.One or more valves may be used to control the flow of fuel through the, or each, hydraulic-to-fuel fuel flow loop. In some implementations, one or more hydraulic-to-fuel fuel pumps may be provided to actively pump fuel to or from the actuators 254; however, in many implementations, the pressure provided by one or more fuel pumps on the main fuel flow path 110 may be sufficient.

[0324] The bleed-off 159 from the main fuel flow path 110 may be located at a different point in the return line 160 of the (or each) fuel flow loop along the main fuel flow path 110. The fuel may be returned to the main fuel flow path 110 upstream, or downstream, of the bleed-off. One or more valves may be used to control the location or locations along the main fuel flow path 110 where the fuel used to actuate the actuators is returned to the main fuel flow path 110. For example, the fuel used for actuation may be returned to a fuel tank 50, 53, just before the combustion chamber 16, or before or after any fuel-oil heat exchanger or fuel pump, as appropriate.Typically, the fuel used for actuation is returned to a point on the main fuel flow path 110 within the engine 10 (as opposed to being returned directly to a fuel tank 50, 53 located elsewhere in the aircraft 1, or to a connecting pipe between the two, for example).

[0325] The supply steps 1010, 1020 may therefore comprise the control of a plurality of fuel flow valves and, possibly also of a plurality of fuel pumps.

[0326] When no actuation is requested, the percentage of fuel redirected to the actuators may be 0%.

[0327] The fuel flow rate to the actuators under idle conditions is generally equal to or greater than the flow rate to the combustion chamber 16 under idle conditions, for example between one and 5.5 times the fuel flow rate to the combustion chamber 16.

[0328] The flow rate to the actuators at cruise is generally less than or similar to the flow rate of the combustion chamber 16, for example between 0.26 times and 1.1 times the fuel flow rate to the combustion chamber 16.

[0329] [Fig. 32] is a flowchart showing a method 1100 of operating a gas turbine engine 10 according to various implementations. The method 1100 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and adapted to be driven by the core shaft 26, the fan comprising 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 powered engine systems 356, 456, 556, 656, 756 including a heat management system 1700, 1900 and a turbine case cooling system 656; and a fuel supply system 152.

[0330] The method 1100 comprises: providing 1110, using the fuel supply system, fuel for combustion in the combustion chamber; and providing 1120, using the fuel supply system, fuel to hydraulically fuel-drive at least three of the plurality of powered engine systems.

[0331] The at least three actuated engine systems may include one or more of: • the VSV 556 system; • the TCC 656 system; • HMU 756; • the 1700 engine heat management system; • the 1900 generator heat management system; and / or • any of the air sampling systems.

[0332] The fuel flow rate to each actuator of the heat management system may be 0 to 15% of the fuel flow rate to the combustion chamber 16 at cruise.

[0333] The fuel flow to each actuator of the heat management system may be 0 to 150% of the fuel flow to the combustion chamber 16 under idle conditions.

[0334] In some implementations, the heat management system may include between 1 and 5 fuel hydraulic actuators.

[0335] The fuel flow to the TCC system may be 0 to 30%, and optionally in the range of 0 to 20%, of the fuel flow to the combustion chamber 16 at cruise.

[0336] The fuel flow to the TCC system may be from 0 to 150%, and optionally in the range of 0 to 75%, of the fuel flow to the combustion chamber 16 at idle conditions.

[0337] As described with respect to method 1000 of [Fig. 31], one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0338] [Fig. 33] is a flowchart showing a method 1200 of operating a gas turbine engine according to various implementations. The method 1200 is executed for an engine 10 comprising:

[0339] an engine core 11 comprising a turbine 19, a combustion chamber 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and adapted to be driven by the core shaft 26, the fan comprising 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 actuators (any or all of actuators 954, 1054, 1154, 1254, which may be more generally referred to as 254), wherein at least one of the plurality of actuators is configured to actuate an air bleed valve 954a, 1054a, 1154a, 1254a; and a fuel delivery system 152.

[0340] The method 1200 comprises: supplying 1210, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and supplying 1220, using the fuel supply system 152, fuel to hydraulically fuel-drive the at least one actuator 954, 1054, 1154, 1254 configured to actuate the air bleed valve 954a, 1054a, 1154a, 1254a.

[0341] The air bleed valve may be any other air bleed valve described above. For example, the air bleed valve may be an anti-icing valve. The air bleed valve may be an upper stage valve. The air bleed valve may be a manifold pressure valve. The air bleed valve may be a high pressure bleed valve, an intermediate pressure valve, or a bypass duct.

[0342] As described with respect to methods 1000, 1100 of Figures 31 and 32, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0343] [Fig. 34] is a flowchart showing a method 1300 of operating a gas turbine engine according to various implementations. The method 1300 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising 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; an engine heat management system 1700; a plurality of actuators, including an actuator 1704, 1704' configured to actuate a valve 1704a, 1704a' in the engine heat management system 1700, wherein the actuator 1704, 1704' configured to actuate a valve 1704a, 1704a' in the engine heat management system 1700 is configured to actuate said valve 1704a, 1704a' so as to allow a non-binary position adjustment between an open valve position and a closed valve position; and a fuel system 152.

[0344] The method 1300 comprises: supplying 1310, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and providing 1320, using the fuel supply system 152, fuel to hydraulically drive by fuel at least the actuator 1704', 1704a' of the engine heat management system valve of the plurality of actuators.

[0345] The engine 10 of some implementations includes multiple actuators 1704, 1704', each configured to actuate a valve 1704a, 1704a' in the engine HMS 1700, and the method 1300 of such implementations may include hydraulically actuating the fuel of multiple such actuators, using the fuel delivery system 152.

[0346] The engine HMS system 1700 may control the temperature of the oil system that lubricates the engine components (e.g., bearing chambers, reduction gear, 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'.

[0347] As described with respect to methods 1000, 1100, 1200 of Figures 31-33, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0348] [Fig. 35] is a flowchart showing a method 1400 of operating a gas turbine engine according to various implementations. The method 1400 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising 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 generator heat management system 1900; a plurality of actuators, including an actuator 1904, 1904' configured to actuate a valve 1904a, 1904a' in the heat management system of generator 1900, wherein the actuator 1904, 1904a' configured to actuate a valve 1904a, 1904a' in the generator heat management system 1900 is configured to actuate said valve 1904a, 1904a' so as to allow non-binary position adjustment between an open valve position and a closed valve position; and a fuel system 152.

[0349] The method 1400 comprises: supplying 1410, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and providing 1420, using the fuel supply system 152, fuel to hydraulically fuel-drive at least the generator heat management system valve actuator 1904, 1904' of the plurality of actuators.

[0350] The engine 10 of some implementations includes multiple actuators 1904, 1904', each configured to actuate a valve 1904a, 1904a' in the generator HMS 1900, and the method 1400 of such implementations may include fuel-hydraulic actuation of multiple such actuators, using the fuel delivery system 152.

[0351] The generator HMS 1900 may be independent of the engine HMS 1700. It may provide cooling for an electrical machine used, for example, to power the airframe with electricity to operate aircraft systems. The generator HMS 1900 includes one or more generator HMS valves 1904a, 1904a' that are controlled by one or more fuel hydraulic actuators 1904, 1904'.

[0352] As described with respect to methods 1000, 1100, 1200, 1300 of Figures 31-34, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0353] It will be understood that the invention is not limited to the implementations described above and that various modifications and improvements may be made without departing from the concepts described herein. Except where mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Claims

1. A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a combustion chamber (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11) and adapted to be driven by the core shaft (26), the fan comprising 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), wherein the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators (254, 554, 654, 754, 854);and a fuel supply system (152), wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber (16), and to supply fuel to hydraulically fuel-drive at least ten of the plurality of actuators (254, 554, 654, 754, 854, 954).;

2. The gas turbine engine (10) of claim 1, wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive at least eleven of the plurality of actuators (254, 554, 654, 754, 854, 954); optionally wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive at least twelve of the plurality of actuators (254, 554, 654, 754, 854, 954).

3. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a variable vane stator system (556), wherein at least two of the plurality of actuators (554a, b) are part of the variable vane stator system (556), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least two of the plurality of actuators (554a, b) that are part of the variable vane stator system (556).

4. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a turbine case cooling system (656), wherein at least two of the plurality of actuators (654a, b) are part of the turbine case cooling system (656), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least two of the plurality of actuators (654a, b) that are part of the turbine case cooling system (656).

5. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a vent valve actuator (854), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel drive the vent valve actuator (854).

6. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises an engine heat management system (1700), and wherein at least one of the plurality of actuators is part of the engine heat management system (1700), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive at least one of the plurality of actuators (1704', 1704a') that is part of the engine heat management system (1700).

7. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a generator heat management system (1900), and wherein at least one of the plurality of actuators is part of the generator heat management system (1900), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least one of the plurality of actuators (1904, 1904') that is part of the generator heat management system (1900).

8. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a hydromechanical unit (756), and wherein at least three of the plurality of actuators are part of the hydromechanical unit, and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least three of the plurality of actuators (754a-c) that are part of the hydromechanical unit (756); optionally, wherein at least four of the plurality of actuators are part of the hydromechanical unit (756), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least four of the plurality of actuators that are part of the hydromechanical unit (756).

9. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a bleed air valve (954a), wherein at least one of the plurality of actuators (254, 954) is configured to actuate the bleed air valve (954a), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive the at least one of the plurality of actuators configured to actuate the bleed air valve (954a).

10. A gas turbine engine (10) according to claim 1, wherein the core shaft (26) directly provides drive to the fan (23), so as to drive the fan at the same rotational speed as the core shaft, such that the engine (10) is a direct drive turbine engine.

11. A gas turbine engine (10) according to claim 1, wherein the turbine engine (10) comprises a reduction gear (30) which receives an input from the core shaft (26) and provides a drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft, such that the engine (10) is a reduction geared turbine engine.

12. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a plurality of actuated engine systems (356, 456, 556, 656, 756, 1700, 1900), including a heat management system (1700) and a turbine case cooling system (656), and wherein the fuel supply system (152) is configured to provide fuel to hydraulically fuel-drive at least three of the plurality of actuated engine systems; optionally wherein the fuel supply system (152) is configured to provide fuel to hydraulically fuel-drive at least five of the plurality of 76 powered engine systems (356, 456, 556, 656, 756, 1700, 1900); or wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive at least seven of the plurality of powered engine systems (356, 456, 556, 656, 756, 1700, 1900).

13. The gas turbine engine (10) of claim 12, wherein at least one of the plurality of actuated engine systems (356, 456, 556, 656, 756, 1700, 1900) comprises at least two of the plurality of actuators (254, 554, 654, 754, 854), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel-drive at least one of the at least two of the plurality of actuators in the at least one actuated engine system.

14. The gas turbine engine (10) of claim 12, wherein at least one of the actuated engine systems (356, 456, 556, 656, 756, 1700, 1900) comprises at least two of the plurality of actuators (254, 554, 654, 754, 854), and wherein the fuel supply system (152) is configured to supply fuel to hydraulically fuel each of the at least two of the plurality of actuators in the at least one actuated engine system.

15. A method (1000) of operating a gas turbine engine (10) for an aircraft, the engine comprising: an engine core (11) comprising a turbine (19), a combustion chamber, a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11) and adapted to be driven by the core shaft (26), the fan comprising 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), wherein the bypass ratio, 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 at cruise conditions, is at least 4; a plurality of actuators (254, 554, 654, 754, 854); and a fuel supply system (152); and wherein the method comprises: providing (1010), using the fuel supply system (152), fuel for combustion in the combustion chamber (16); and providing (1020), using the fuel supply system (152), fuel to hydraulically fuel-drive at least ten of the plurality of actuators (254, 554, 654, 754, 854, 954).

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