A power distribution system for an aircraft, and method of operating includes, including determining, by a supplemental electrical power system, an operating mode of the aircraft, selecting, by the supplemental electrical power system, at least one supplementary power source of a set of at least two supplementary power sources, and based on the operating mode of the aircraft at least one of receiving, by the selected at least one supplementary power source, a first current from the primary power bus, or selectively providing a second current to the primary power bus from the selected at least one supplementary power source.
H02J 7/00 - Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
H02J 7/14 - Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries for charging batteries from dynamo-electric generators driven at varying speed, e.g. on vehicle
H02M 3/335 - Conversion of DC power input into DC power output with intermediate conversion into AC by static converters using discharge tubes with control electrode or semiconductor devices with control electrode to produce the intermediate AC using devices of a triode or a transistor type requiring continuous application of a control signal using semiconductor devices only
2.
ENERGY MANAGEMENT SYSTEM AND METHOD OF OPERATING FOR A FLEET OF AIRCRAFT
An energy management system and method of operating the energy management, which include estimating an energy demand for flight plans for a fleet of aircraft. The flight plans are received from a flight plan database. The system is configured to determine whether a set of dischargeable energy modules are locatable at a respective location of a subset of the fleet of aircraft based at least in part on a replaceable power source inventory database or the subset of the plurality of flight plans. The system is configured to generate a power source inventory distribution plan allocating a subset of dischargeable energy modules for the subset of the plurality of flight plans for the fleet of aircraft based at least in part on the determination that the set of dischargeable energy modules are locatable at the respective location of the subset of the fleet of aircraft.
H02J 7/00 - Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
B60L 50/50 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells
B60L 58/18 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries of two or more battery modules
B64D 27/355 - Arrangements for on-board electric energy production, distribution, recovery or storage using fuel cells
B64D 27/357 - Arrangements for on-board electric energy production, distribution, recovery or storage using batteries
B64D 27/359 - Arrangements for on-board electric energy production, distribution, recovery or storage using capacitors
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
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A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
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NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
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AXIAL
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A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
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FT
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AXIAL
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NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
A turbine engine for an aircraft includes a fuel delivery assembly for a hydrocarbon fuel to flow therethrough, a combustor combusting the fuel to generate combustion gases, and a core air exhaust nozzle exhausting the combustion gases from the turbine engine. The turbine engine also includes a contrail mitigation system having a heater and a fuel precipitate separator. The heater is selectively operable to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel, and the fuel precipitate separator separates the fuel precipitates generated by the heater from the fuel. A controller is coupled to the heater to operate the heater to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel in response to a contrail mitigation input.
A heat exchanger includes an outer body defining a centerline axis and a flow passage extending from an inlet to an outlet of the heat exchanger. A first annular fin is disposed within the flow passage, wherein the first annular fin is concentrically aligned with the centerline axis and defines a first undulating surface. A second annular fin is disposed within the flow passage. The second annular fin is concentrically aligned with the centerline axis and defines a second undulating surface. The second annular fin is radially spaced from the first annular fin to define a flow channel therebetween. The flow passage, the first annular fin, and the second annular fin diverge along the centerline axis downstream from the inlet.
A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
F02K 3/065 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front and aft fans
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
8.
TURBINE ENGINE WITH COMBUSTION SECTION AND FUEL PASSAGE THAT SUPPLIES HYDROGEN-CONTAINING FUEL TO COMBUSTION SECTION
A turbine engine comprising a compression section, combustion section, and turbine section is serial flow arrangement, with the combustion section including an injector for providing a mixture of fuel and air for combustion. The injector includes a body, an inner nozzle provided within the body and defining an injector axis, and an outer nozzle in annular arrangement about the inner nozzle. A first fuel passage fluidly couples to the inner nozzle and a second fuel passage fluidly couples to the outer nozzle. A first set of air conduits are in annular arrangement about the body interior of the outer nozzle and a second set of air conduits are in annular arrangement about the body exterior of the outer nozzle.
An aircraft is provided defining a longitudinal direction, a lateral direction, and a longitudinal centerline extending along the longitudinal direction, the aircraft including: a body; a pair of wings extending outward from the body along the lateral direction, each wing of the pair of wings defining a leading edge; and a propulsion system comprising a first engine and a second engine spaced from one another along the lateral direction, the propulsion system further comprising a first thrust vectoring system operable with the first engine and a second thrust vectoring system operable with the second engine.
A gas turbine engine is provided, comprising: a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order; and a fan section having a fan drivingly coupled to the turbomachine and an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine, the airflow surface defining a plurality of boundary layer openings configured to ingest a boundary layer of the fan airflow over the airflow surface during operation of the gas turbine engine.
An insertion tool with an implement gripping mechanism is provided. The insertion tool includes an elongated section defining an implement path extending from a proximal end to a distal end, a handle coupled to the proximal end of the elongated section, a grip mechanism coupled to the handle and configured to selectively secure an implement inserted into the implement path through the handle, and a grip actuator operable to cause the grip mechanism to secure the implement to the handle.
A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.
Sensor information from a plurality of sensors that monitor a heat exchanger is accessed and then sensed information that corresponds to that sensor information is input into a control circuit configured to output virtual parameters corresponding to the heat exchanger as a function, at least in part, of the sensed information. A particular fault class is determined from amongst a plurality of fault classes as a function, at least in part, of the virtual parameters and at least one task is identified regarding the heat exchanger that corresponds to the particular fault class.
B64D 13/08 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space the air being conditioned the air being heated or cooled
A propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine including a turbo-engine, a first propulsor including a first propulsor shaft, and a first gearbox assembly having a gear assembly. The propulsion system includes a second propulsor that is remote from the turbine engine and includes a second gearbox assembly including a gear assembly. The propulsion system further includes a lubrication system including a first gearbox lubrication system, an engine lubrication system, and a second gearbox lubrication system. The first gearbox lubrication system is disposed within the first gearbox assembly and supplies lubricant to the gear assembly. The engine lubrication system supplies lubricant to the engine bearings. The first gearbox lubrication system is fluidly separate from the engine lubrication system. The second gearbox lubrication system is disposed within the second gearbox assembly and supplies lubricant to the gear assembly.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
B64C 11/48 - Units of two or more coaxial propellers
B64D 35/02 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants
B64D 35/04 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion, which can supply a primary fuel supply and a secondary fuel supply. Increasing efficiency and reducing emission require the use of alternative fuels, which combust at higher temperatures or burn at faster burn speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.
A fan assembly for a gas turbine engine includes a fan actuation system, a fan blade hub, a plurality of fan blades, and a counterweight system arranged disconnected from the plurality of fan blades. The counterweight system includes a counterweight hub, and a plurality of counterweight levers each having a counterweight trunnion rotationally connected to the counterweight hub, a cantilever arm with a counterweight connected thereto. The fan blades are rotationally connected to the fan blade hub and to the fan actuation system, and the counterweight levers are rotationally connected to the counterweight hub and to the fan actuation system. The fan actuation system is arranged to correspondingly rotate each of the plurality of fan blades about a respective fan blade pitch change axis, and each of the plurality of counterweight levers about a counterweight lever rotational axis in unison.
A turbofan engine for an aircraft includes a core cowl, a nacelle assembly positioned radially outward of the core cowl defining a bypass airflow passage between the core cowl and the nacelle assembly where the bypass airflow passage has a fan exit nozzle. The nacelle assembly includes a fan cowl, a transcowl positioned aft of the fan cowl, and a thrust reverser assembly. An actuation assembly is operably connected to at least one of the transcowl or the thrust reverser assembly and is actuatable to move the transcowl aft from a first position where the cascade assembly is covered to a second position where the cascade assembly is uncovered. The actuation assembly is further actuatable to move the transcowl forward from the first position to a third position to reduce an area of the fan exit nozzle.
F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
A turbine engine comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising a combustion chamber, a fuel nozzle assembly comprising a fuel supply passage and having a fuel supply passage outlet fluidly coupled to a combustion chamber, and an optical sensor located in the fuel supply passage and oriented to sense a combustion flame in the combustion chamber.
F23N 5/08 - Systems for controlling combustion using devices responsive to thermal changes or to thermal expansion of a medium using light-sensitive elements
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
21.
Control Logic for Thrust Link Whiffle-Tree Hinge Positioning for Improved Clearances
Systems and methods for optimizing clearances within an engine include an adjustable coupling configured to couple a thrust link to the aircraft engine, an actuator coupled to the adjustable coupling, where motion produced by the actuator adjusts a hinge point of the adjustable coupling, sensors configured to capture real time flight data, and an electronic control unit. The electronic control unit receives flight data from the sensors, implements a machine learning model trained to predict clearance values within the engine based on the received flight data, predicts, with the machine learning model, the clearance values within the engine based on the received flight data, determines an actuator position based on the clearance values, and causes the actuator to adjust to the determined actuator position.
F01D 11/14 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
F01D 11/22 - Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
F01D 11/24 - Actively adjusting tip-clearance by selectively cooling or heating stator or rotor components
F02C 7/20 - Mounting or supporting of plantAccommodating heat expansion or creep
Sensor information from a plurality of sensors that monitor an apparatus is accessed and then sensed information that corresponds to that sensor information is input into a control circuit configured to output virtual parameters corresponding to the apparatus as a function, at least in part, of the sensed information. A particular fault class is determined from amongst a plurality of fault classes as a function, at least in part, of the virtual parameters and at least one task is identified regarding the apparatus that corresponds to the particular fault class.
An airfoil for a gas turbine engine. The airfoil includes an inner portion including a hollow fiber composite, and a surface portion including a solid fiber composite. The surface portion at least partially surrounds the inner portion. The hollow fiber composite of the inner portion can be a woven composite with at least one of a warp fiber or a weft fiber including a hollow carbon fiber. The hollow fiber composite of the inner portion also can be a braided composite with at least one strand including a hollow carbon fiber.
A propulsion system and a method of operating the propulsion system. The propulsion system includes a first turbine engine, a second turbine engine, and a lubrication system. The lubrication system supplies a lubricant to the first turbine engine and the second turbine engine. The lubrication system includes one or more first lubricant sensors that sense information about the lubricant to the first turbine engine and one or more second lubricant sensors that sense information about the lubricant to the second turbine engine. The first controller and the second controller receive the information of the lubricant from the first lubricant sensors and the second lubricant sensors. The first controller controls the lubrication system to supply the lubricant to the second turbine engine based on the sensed information about the lubricant in the second turbine engine received at the first controller when the second turbine engine is shut down.
A propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine, an electric power supply, and a lubrication system. The turbine engine includes a turbo-engine having a low-pressure shaft and one or more engine bearings, a propulsor having a propulsor shaft, a gearbox assembly having a gear assembly, and an electric machine. The propulsor shaft is drivingly coupled to the low-pressure shaft through the gear assembly. The electric machine powers the propulsor when the turbo-engine is shut down. The lubrication system supplies a lubricant to at least one of the one or more engine bearings or the gear assembly. The electric power supply powers the lubrication system when the turbo-engine is shut down.
A turbine engine control system for a propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine including a turbo-engine and a first propulsor drivingly coupled to the turbo-engine, and a second propulsor that is remote from the turbine engine. The turbine engine control system includes a single throttle lever and a controller that receives an input from the single throttle lever, and controls the turbine engine and the second propulsor based on the input from the single throttle lever.
A turbine engine includes a turbo-engine, a fan, a frame, and a lubrication system. The turbo-engine includes a core air flowpath. The fan is drivingly coupled to the turbo-engine. The frame supports the core air flowpath and includes a plurality of lubricant struts that extend through the core air flowpath. The lubrication system includes a sump having a lubricant therein, a scavenge reservoir, and a lubricant strut flowpath disposed through each of the plurality of lubricant struts. The lubricant strut flowpath is in fluid communication with the sump and the scavenge reservoir. The lubricant strut flowpath of each of the plurality of lubricant struts directs the lubricant from the sump to the scavenge reservoir.
An airfoil assembly for a turbine engine, the airfoil assembly including a platform defining an inner surface and an outer surface, a variable pitch airfoil extending radially from the outer surface of the platform from a root to a tip to define a span length and a mounting structure connected to the platform.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor. The unducted fan propulsors can also include a first VPF parameter within a range of 0.10 to 0.40 and defined as the hub-to-tip radius ratio divided by the fan pressure ratio and/or a second VPF parameter within a range of 1-30 lbf/in2 and defined as the bearing spanwise force divided by the fan area. In certain examples, the unducted fan propulsor further includes a pitch change mechanism and/or a gearbox.
B64D 27/18 - Aircraft characterised by the type or position of power plants of jet type within, or attached to, wings
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
30.
AUTOMATED DE-POWDERING OF ADDITIVE MANUFACTURING BUILD
An automated de-powdering system comprises a sleeve alignable with at least a portion of a build chamber and at least one actuator actuatable with respect to a build plate to move at least a portion of an additive manufacturing build out of the build chamber and into the sleeve. At least one support assembly is couplable to the sleeve and includes at least one support member insertable through the sleeve and into the additive manufacturing build. At least one agitation mechanism is couplable to at least one of the at least one support assembly or the sleeve and is actuatable to at least partially convey the powder build material away from at least one object of the additive manufacturing build. The support member is positioned to support the object in an absence of the powder build material around the object.
B29C 64/153 - Processes of additive manufacturing using only solid materials using layers of powder being selectively joined, e.g. by selective laser sintering or melting
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
A propulsor assembly for a turbine engine includes a propulsor and a propulsor actuation system. The propulsor has a plurality of propulsor blades. Each of the plurality of propulsor blades is rotatable about a blade pitch axis. The propulsor actuation system includes an actuator for rotating the plurality of propulsor blades and a trunnion mechanism that includes a plurality of trunnion assemblies. Each of the plurality of trunnion assemblies is coupled to a respective one of the plurality of propulsor blades and includes an outer sleeve coupled to a blade spar of the respective propulsor blade, an inner sleeve coupled to the outer sleeve, and an actuation member that engages the inner sleeve. The actuation member is rotatably engageable by the actuator to rotate the respective propulsor blade about the blade pitch axis.
An additive manufacturing testing apparatus including an interchangeable base plate, an interchangeable dolly, and a sensor. The interchangeable base plate includes a transparent section over which a photocurable material is disposed. The interchangeable dolly is positionable between an upper position and a lower position along a vertical axis. The interchangeable dolly is configured to compress the photocurable material between the transparent section and the interchangeable dolly and configured to retract from the transparent section. The sensor is coupled to the interchangeable dolly. The sensor is configured to measure tensile load and compressive load during movement of the interchangeable dolly between the upper position and the lower position along the vertical axis during compression of the photocurable material and retraction of the interchangeable dolly from the transparent section.
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain that is coupled in gear with the low pressure spool. The low pressure geartrain is coupled in gear with the low pressure spool and the high pressure geartrain is coupled in gear with the high pressure spool at a common axial location relative to a centerline axis of the gas turbine engine.
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain coupled in gear with the low pressure spool. The gas turbine engine further includes a power transfer device coupling the high pressure geartrain and the low pressure geartrain to transfer power between the high pressure spool and the low pressure spool.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
35.
GAS TURBINE ENGINE HAVING A MECHANICAL POWER SHARING ARRANGEMENT
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain that is coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain that is coupled in gear with the low pressure spool. The gas turbine engine further includes an accessory gearbox that has an accessory drive gear that is selectively drivingly coupled at least one of the high pressure geartrain and the low pressure geartrain.
A blended wing aircraft is provided, including: an aircraft engine comprising a lubrication system; a body having a fuselage and a pair of wings extending outward from the fuselage; and an access panel assembly operable with the lubrication system of the aircraft engine, the access panel assembly positioned on or within the body at a location remote from the aircraft engine.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
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A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
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NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F04D 19/00 - Axial-flow pumps specially adapted for elastic fluids
38.
GEARBOX ASSEMBLY HAVING A SEGMENTED SCAVENGE GUTTER FOR A TURBINE ENGINE
A gearbox assembly for a turbine engine. The gearbox assembly includes a gearbox having a plurality of gears, and a scavenge gutter radially outward of the gearbox and mounted to a fan frame flange of a fan frame of the turbine engine, the scavenge gutter being configured to collect lubricant that is ejected by the plurality of gears of the gearbox during rotation of the plurality of gears. An interface between the fan frame flange and the scavenge gutter is located at a first radial distance R1 relative to a longitudinal centerline axis of the turbine engine and the scavenge gutter is located at a second radial distance R2 relative to the longitudinal centerline axis. The second radial distance R2 is greater than the first radial distance R1. The scavenge gutter is segmented and includes a plurality of sector portions joined together to form the scavenge gutter.
A turbine engine includes a fan having a plurality of fan blades, a nacelle that surrounds the fan, and a fan blade sensor system. The fan blade sensor system includes a plurality of fan blade sensors disposed in the nacelle to sense the plurality of fan blades as the plurality of fan blades rotates. The fan blade sensor system also includes a controller that determines a fan speed of the fan based on a rotational time between the fan blades as sensed by the plurality of fan blade sensors.
Systems, apparatus, computer-readable medium, and associated methods for secure additive manufacturing are disclosed. An example apparatus includes an inbound one-way data diode to receive, authenticate, and route an inbound file in a first direction within a secure additive manufacturing system, the inbound one-way data diode unable to transmit data out of the secure additive manufacturing system in a second direction. The example apparatus includes an additive manufacturing machine to build a part, the build of the part adjusted by the inbound file when authenticated by the inbound one-way data diode. The example apparatus includes an outbound one-way data diode to authenticate and transmit outbound data in the second direction to an external system outside the secure additive manufacturing system, the outbound one-way data diode unable to transmit data into the secure additive manufacturing system in the first direction.
G06F 21/62 - Protecting access to data via a platform, e.g. using keys or access control rules
B29C 64/393 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
B33Y 50/02 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
G06F 21/57 - Certifying or maintaining trusted computer platforms, e.g. secure boots or power-downs, version controls, system software checks, secure updates or assessing vulnerabilities
A lubrication system for a turbine engine that includes one or more rotating components. The lubrication system includes one or more tanks that store lubricant, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during stable operating conditions of the lubrication system. The auxiliary lubrication system includes an auxiliary feed line and an auxiliary supply line. The auxiliary lubrication system receives the lubricant from the one or more tanks through the auxiliary feed line. The auxiliary lubrication system supplies the lubricant to the one or more rotating components through the auxiliary supply line when there is a potential lubricant interruption in the lubrication system.
A thermal management system for a propulsion system of an aircraft includes a fluid circuit configured to provide a flow of a compressed fluid from a compressor section of the propulsion system to, in serial flow order, a first turbine, a compressor, a second turbine, a thermal load, and an exhaust sink. A heat exchanger is disposed in a fan stream of the propulsion system at a location along the fluid circuit between the compressor and the second turbine. The fluid circuit extends through and is in thermal communication with the heat exchanger.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
⨯
D
FT
L
AXIAL
⨯
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
⨯
D
FT
L
AXIAL
⨯
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F04D 19/00 - Axial-flow pumps specially adapted for elastic fluids
A pulse illumination imaging system is provided. The system includes an image sensor, a light source, and a controller. The image sensor includes a plurality of light sensitive pixel elements that are activatable for a designated exposure time to capture one or more images. The controller is configured to determine an activation time to activate the image sensor; activate the image sensor at the activation time; and activate the light source during the exposure time of the image sensor to produce a pulse having a preconfigured time duration that is less than the exposure time of the image sensor.
A fuel tank heat rejection system for an aircraft. The fuel tank heat rejection system includes a fuel tank compartment in the aircraft and a fuel tank having an exterior surface and storing fuel therein. The fuel tank is located in the fuel tank compartment. The fuel tank heat rejection system includes one or more air valves that provide fluid communication to the fuel tank compartment. The one or more air valves opening to operably direct cooling air into the fuel tank compartment through the one or more air valves, the cooling air contacting the exterior surface of the fuel tank such that heat from the fuel is rejected from the fuel tank.
A gas turbine engine defining an axial direction and a radial direction includes a spinner defining a spinner duct and a spinner inlet to the spinner duct, a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a fan duct, a fan duct inlet to the fan duct, a core duct, and a core inlet to the core duct, and a primary fan driven by the turbomachine, wherein the spinner inlet is upstream of the primary fan.
An example booster splitter includes an inner cylindrical structure; an outer cylindrical structure concentric with the inner cylindrical structure; an annular lip having an arc extending between the inner cylindrical structure and the outer cylindrical structure; a first protrusion at a first circumferential position on the annular lip, the first protrusion protruding axially from the annular lip and extending from the outer cylindrical structure to the inner cylindrical structure along the arc of the annular lip; and a second protrusion at a second circumferential position on the annular lip.
F01D 25/02 - De-icing means for engines having icing phenomena
B64D 15/00 - De-icing or preventing icing on exterior surfaces of aircraft
B64D 15/16 - De-icing or preventing icing on exterior surfaces of aircraft by mechanical means, e.g. pulsating mats or shoes attached to, or built into, surface
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
A heat exchanger assembly includes a manifold, a plurality of plates supported by the manifold, a bypass channel in fluid communication with the manifold, and a flow controller fluidly connecting a heated fluid supply to the bypass channel. The flow controller is configured to flow a heated fluid from the heated fluid supply through the bypass channel.
F02K 3/065 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front and aft fans
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
51.
METHOD OF SUPPLYING A FUEL TO A COMBUSTOR FOR A TURBINE ENGINE
A turbine engine includes a compression section, combustion section, and turbine section is serial flow arrangement. A supply of gaseous fuel is mixed with a supply of inert gas to form a mixture of fuel and inert gas. The mixture of fuel and inert gas is provided to or intermixed within a fuel injector for combustion within a combustor provided within the combustor section to drive the turbine section.
Truss-braced wing aircraft engine mounting and associated systems are disclosed. An example aircraft includes a fuselage including a central structural section, a wing extending from the fuselage, a truss extending from the central structural section and coupled to the wing, a pylon coupled to the central structural section, and an engine coupled to the pylon, wherein the pylon is positioned at an angle substantially 45 degrees from a horizontal plane and a vertical plane.
A combustor having a plurality of openings to control a vortex driver jet within the combustor. The combustor includes an outer casing and an inner casing extending circumferentially about a longitudinal combustor centerline axis, an outer liner spaced apart from the outer casing to define therebetween an outer flow passage and an inner liner spaced apart from the inner casing to define therebetween an inner flow passage, a dome structure, the outer liner and the inner liner defining a combustion chamber, a plurality of outer openings provided in the outer casing and a plurality of inner openings provided in the inner casing. The plurality of outer openings and inner openings are configured to bleed airflow from or to introduce airflow into the outer flow passage or the inner flow passage, to control the vortex driver jet within the driver openings, to drive a vortex in the combustion chamber.
GE Marmara Technology Center Muhendislik Hizmetleri Ltd (Turkey)
GE Aerospace Poland Sp. z o.o. (Poland)
Inventor
Unsal, Arda
Whitener, Geoffrey
Bibler, John David
Yilmaz, Batu
Beyer, Katherine
Pazinski, Adam Tomasz
Abstract
An engine includes a compressor including an inner casing and an outer casing where the inner casing defines a primary flow path for a primary airflow through the compressor. The inner casing and the outer casing define a bleed air cavity therebetween. The inner casing at least partially defines a bleed air channel extending circumferentially about the inner casing to direct a bleed airflow from the primary airflow into the bleed air cavity. One or more flow control devices located circumferentially about the compressor to actively or passively circumferentially balance a flow of the bleed airflow into the bleed air cavity or within the bleed air cavity, wherein the one or more flow control devices are disposed at least partially within the bleed air channel, form at least part of the bleed air channel, or extend axially aft from the bleed air channel.
A binder solution comprises a fugitive metal precursor, a thermoplastic binder, and a solvent. The fugitive metal precursor may comprise an alkaline earth metal, a transition metal, a post-transition metal, a metalloid, a rare earth metal, or combinations thereof. The fugitive metal precursor may comprise a salt such as carboxylate, nitrate, sulfate, carbonate, formate, chloride, halide, derivatives thereof, and combinations thereof. A method of manufacturing a part includes depositing a layer of particulate material on a working surface, selectively applying a binder solution into the layer of particulate material in a pattern representative of a layer of the part, repeating the steps of depositing and selectively applying to form a plurality of layers of particulate material with the applied binder solution, and curing the applied binder solution in the plurality of layers of particulate material with the applied binder solution to evaporate the solvent and form a green body part.
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Li, Cong
Yi, Xuan
Younsi, Karim
Huang, Shenyan
Xiong, Han
Hanczewski, Pawel Piotr
Klausen, Michael
Abstract
An electrical cable assembly includes a cable portion and a connector housing portion coupled to the cable portion. The cable portion includes a first electrical conductor circumferentially surrounded by a first set of layers. The first set of layers includes a first semiconductive layer, a first insulative layer, a second semiconductive layer, a magnetic layer, a second insulative layer, and an electrically conductive. The connector housing portion includes a connector housing, a printed circuit board disposed within the connector housing; and a set of capacitors mounted to the printed circuit board and coupled to the magnetic layer.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F01D 7/02 - Rotors with blades adjustable in operationControl thereof having adjustment responsive to speed
F01D 17/26 - Devices dealing with sensing elements or final actuators or transmitting means between them, e.g. power-assisted the operation or power assistance being predominantly non-mechanical fluid, e.g. hydraulic
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F03D 7/02 - Controlling wind motors the wind motors having rotation axis substantially parallel to the air flow entering the rotor
F03D 17/00 - Monitoring or testing of wind motors, e.g. diagnostics
A gas turbine engine includes a turbomachine comprising a low pressure (LP) spool and a high pressure (HP) spool that rotate about a central axis, an electric motor mechanically coupled to the LP spool for selectively rotating the LP spool, a starter assembly mechanically coupled to the HIP spool for selectively rotating the HP spool, and a controller in operative communication with the electric motor and the starter assembly, the controller being configured to operate the electric motor to rotate the LP spool and operate the starter assembly to rotate the HIP spool during engine startup.
A turbine component comprised of a titanium alloy that has been modified from Ti-64 is provided. The modification preserves the desired properties of Ti-64 (e.g., relatively isotropic properties, a relatively low density, tolerance to FOD, repairability, and low cost) while improving the thick section strength, HCF capability, creep strength, and low deformation following FOD to approach those beneficial aspects of Ti-17 and Ti-6246. Methods of forming such turbine components are also provided.
Methods of forming a coating are presented. For example, a method for forming a coating on particles may include mixing an initial powder with a pressing medium, the initial powder comprising a plurality of core particles having an initial shell coating thereon; isostatic pressing the initial powder within the pressing medium to densify the initial shell coating on the plurality of core particles to form a pressed powder, the pressed powder comprising a densified shell coating on the plurality of core particles; and thereafter, removing the pressing medium from the pressed powder.
Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 6/18 - Plural gas-turbine plantsCombinations of gas-turbine plants with other apparatusAdaptations of gas-turbine plants for special use using the waste heat of gas-turbine plants outside the plants themselves, e.g. gas-turbine power heat plants
F02C 7/10 - Heating air supply before combustion, e.g. by exhaust gases by means of regenerative heat-exchangers
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. Increasing efficiency and meeting emission needs can be met with the use of alternative fuels, which combust at higher temperatures or higher speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.
A combustor liner for a combustor of a gas turbine includes an outer liner having a plurality of outer liner segments, and an inner liner having a plurality of inner liner segments. Each segment of the inner liner and the outer liner includes at least one slotted dilution opening therethrough extending in a circumferential direction, and each slotted dilution opening includes a deflector wall extending radially from the respective liner into a dilution zone of a combustion chamber between the outer liner and the inner liner. The at least one slotted dilution opening may be a curved slot (either concave or convex) dilution opening, and the curved slot dilution openings for each segment of the outer liner and the inner liner may be connected so as to provide a wavy slotted dilution opening extending annularly through the outer liner and the inner liner.
An apparatus for maintaining a gas turbine engine having at least one port comprises a tool having an end effector to effect maintaining the gas turbine engine, the tool being configured to temporarily enter and exit the gas turbine engine via the at least one port and having a first portion and a second portion that are separated by at least a first area of articulation. A first inertial measurement unit is affixed with respect to that first portion and a second inertial measurement unit is affixed with respect to that second portion. A control circuit operably couples to those inertial measurement units and receives corresponding information regarding those portions of the tool. The control circuit can then process that received information to generate positional proprioception information as regards those monitored tool portions.
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
A gas turbine engine includes a fan, a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining in part a working gas flowpath, the compressor section including a frame and defining a cavity forward of the frame, the compressor section further comprising a booster disposed in the working gas flowpath including a vane in fluid communication with the working gas flowpath and the cavity, the gas turbine engine defining a bypass passage over the turbomachine, and a heat exchanger disposed at least partially in the cavity, wherein the heat exchanger is in fluid communication with the booster in the working gas flowpath and is in fluid communication with the bypass passage.
An aerodynamic device defining a thickness direction is provided. The aerodynamic device configured to produce lift or thrust or configured to be a part of an aerodynamic system that produces lift or thrust. The aerodynamic device includes a cowl assembly that defines at least in part an airflow stream. The cowl assembly includes a first cowl and a second cowl moveable relative to the first cowl. The first cowl includes a plurality of first cowl indentations at an end of the first cowl. The second cowl defines an outer surface along the thickness direction and an inner surface along the radial direction. The second cowl includes a plurality of second cowl indentations complementary in shape to the plurality of first cowl indentations. The plurality of second cowl indentations are positioned locally on the outer surface of the second cowl or locally on the inner surface of the second cowl.
A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.
An analog system for edge artificial intelligence computing includes a first plurality of analog edge devices configured to receive an input analog signal and to output a first plurality of output analog signals, a second plurality of analog edge devices configured to receive the first plurality of output analog signals and to output a second plurality of output analog signal, and one or more memory devices in communication with the first plurality of analog edge devices and the second plurality of analog edge devices, and configured to store weight parameters, the weight parameters being adjustable based on time constants of the first plurality of analog edge devices or the second plurality of analog devices, or both. The second plurality of output analog signals are multiplied by the weight parameters to obtain a plurality of weighted analog signals.
Heat exchangers for gas turbine engines are described herein. An example heat exchanger includes a first plenum, a second plenum, and divider plates coupled to and extending between the first and second plenums. The divider plates are spaced apart from each other. Each of the divider plates has one or more internal passages for a fluid to flow between the first and second plenums. The heat exchanger also includes stiffeners between each adjacent pair of the divider plates, wherein at least one of the stiffeners is skewed relative to a connecting one of the divider plates.
A gas turbine engine includes a fan assembly, a turbo-engine encased within a turbo-engine cowl structure, a nacelle, and a thrust reverser system arranged, at least in part, in the nacelle. The turbo-engine includes a low-pressure compressor, a high-pressure compressor, an inter-compressor frame structure between the low-pressure compressor and the high-pressure compressor and including an outer frame portion having a cowl door engagement member on an upstream side of the outer frame portion, and a plurality of cowl doors defining the turbo-engine cowl structure. Each cowl door includes an inter-compressor frame engagement portion that engages with the cowl door engagement member of the inter-compressor frame structure, and a plurality of thrust reverser drag link connectors arranged on an outer side of the cowl door and arranged between the upstream side of the inter-compressor frame structure and a downstream side of the inter-compressor frame structure.
F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
F02C 9/52 - Control of fuel supply conjointly with another control of the plant with control of working fluid flow by bleeding or by-passing the working fluid
A gas turbine engine including a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section having a fuel nozzle including a fuel nozzle body defining an axis and having an inner surface defining a channel fluidly coupled to a combustion chamber, a support matrix located within the channel comprising a plurality of segments which intersect each other when viewed from aft, and a set of vortex generators located on the support matrix.
A combustor including a dome structure, an inner liner and an outer liner connected to the dome structure to define a combustion chamber, and a first segment coupled to the outer liner and a second segment coupled to the inner liner, the first segment including a first geometric ramp and the second segment including a second geometric ramp. The first geometric ramp and the second geometric ramp have one or more driver holes, driver slots, and/or a plurality of driver vanes. An upstream crossflow enters the one or more driver holes, driver slots, and/or the plurality of driver vanes to generate an airflow jet having an increased angle at an exit of the one or more to driver holes, driver slots, and/or the plurality of driver vanes relative to a surface of the inner liner or a surface of the outer liner.
A system may include an image sensor. A system may include an actuator configured to cause a controlled movement of the image sensor relative to a target element, the controlled movement being based on an operating velocity of the target element relative to an initial position of the image sensor. A system may include a controller communicatively coupled to the image sensor, the controller configured to: identify the operating velocity, determine an activation time to activate the image sensor for a designated exposure time based on the operating velocity and the controlled movement; and activate the image sensor at the activation time.
A gas turbine engine includes a turbomachine, a primary fan driven by the turbomachine, a secondary fan, a booster, and an outlet guide vane. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The secondary fan is located downstream of the primary fan within the inlet duct. The booster is located downstream of the secondary fan and includes a booster rotor blade and booster cowl that separates an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet. The outlet guide vane is positioned downstream of the secondary fan and upstream of the upper fan duct inlet or positioned within the upper fan duct.
A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, a booster located downstream of the secondary fan and comprising a booster rotor blade and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, and a flow blocker located at the lower fan duct inlet and movable from an open position to a closed position, wherein, in the closed position, the flow blocker blocks a flow through at least a portion of the lower fan duct inlet.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
F02C 3/00 - Gas-turbine plants characterised by the use of combustion products as the working fluid
A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, and a booster located downstream of the secondary fan and comprising a booster rotor blade, an inlet guide vane, and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet, the inlet guide vane located forward of the booster rotor blade.
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, a core cowl, and a booster located downstream of the secondary fan and including a booster rotor blade and a booster cowl, the booster cowl located outward of the booster rotor blade and within the fan duct at the fan duct inlet, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the booster including a midspan shroud coupled to the booster rotor blade.
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F01D 5/22 - Blade-to-blade connections, e.g. by shrouding
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
78.
MAPPING BETWEEN TIME-SENSITIVE NETWORK DATA PACKETS AND ARINC 664 PART 7 DATA PACKETS
Systems, methods, and other embodiments described herein relate to mapping data packets between an ARINC 664 Part 7 (A664P7) data format and a Time-Sensitive Network (TSN) data format. In one embodiment, a method includes, in response to receiving a first data packet in a first data transmission format from a first data end system, forming a second data packet in a second data transmission format based on the first data packet. The first data end system is capable of transmitting and receiving data packets in the first data transmission format. The first data transmission format is one of A664P7 data format and TSN data format and the second data transmission format is the other of A664P7 data format and TSN data format. The method further includes transmitting the second data packet to a second data end system that is capable of receiving data packets in the second data transmission format.
H04L 69/325 - Intralayer communication protocols among peer entities or protocol data unit [PDU] definitions in the network layer [OSI layer 3], e.g. X.25
A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gear includes a first gear that is a split sun gear including a forward sun gear and an aft sun gear separate from the forward sun gear. The forward sun gear and the aft sun gear are each rotationally coupled to a rotating shaft of the gas turbine engine.
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. Varying the geometry of the swirler can provide for improved supply of air, which can improve efficiency and flame control.
A diffuser in flow communication with a combustor of a turbomachine engine. The diffuser includes an outer annular wall, an inner annular wall, and at least one tube extending from one of an outer passlet defined by the outer annular wall or an inner passlet defined by the inner annular wall. The inner annular wall and the outer annular wall together define a primary passage therebetween that extends in an aft direction of the diffuser such that air flows through the primary passage in the aft direction. The at least one tube is arranged such that a portion of the air within the primary passage flows through the at least one tube at an angle relative to the aft direction.
A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius RFan_LE and a trailing edge fan radius RFan_TE, and the fan defining a leading edge hub radius RHub_LE and a trailing edge hub radius RHub_TE, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.
There are provided methods and systems for making or repairing a specified part. For example, there is provided a method for creating an optimized manufacturing process to make or repair the specified part. The method includes receiving by a system configured to make or repair the specified part and from a machine communicatively coupled with the system, a set of sensor data and/or inspection data associated with at least one of an additive and a reductive manufacturing or repair process or with at least one of a pre-treatment and a post-treatment step. The method includes creating an optimized manufacturing process to make or repair the specified part, the creating including. The method includes updating, in real time, a surrogate model corresponding with a physics-based model of the specified part, wherein the surrogate model forms a digital twin of the specified part. The method includes further updating the surrogate model with the sensor data and/or inspection data. The method includes executing, based on the digital twin, the optimized manufacturing process to either repair or make the specified part.
B29C 64/393 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
B29C 64/171 - Processes of additive manufacturing specially adapted for manufacturing multiple 3D objects
B29C 64/188 - Processes of additive manufacturing involving additional operations performed on the added layers, e.g. smoothing, grinding or thickness control
A gas turbine engine includes a turbomachine including a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a fan duct inlet to a fan duct, and a core inlet to a core duct, a fan driven by the turbomachine, and a booster downstream of the fan, the booster comprising a booster rotor blade and a booster cowl.
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
85.
METHOD AND SYSTEM FOR A BATTERY MONITORING CIRCUIT
A method and system for determining the health of a set of batteries through the use of a battery monitoring circuit. The battery monitoring circuit including a first current loop and a second current loop. The first current loop being enabled by a first switch, a first resistor and a second switch. The second current loop being enabled by the first switch, a third switch, a voltage sensor, and the second switch.
A variable pitch fan for a gas turbine engine includes a hub, a fan blade supported by the hub, the fan blade including a blade root, and a retention system including a bearing and a retainer, the retainer securing the bearing to the blade root. The hub is rotatably supported by the bearing.
An engine for aeromechanical instability abatement includes a sensor configured to capture data from rotating blades of the engine system, a airflow effector device, and an engine controller. The engine controller is configured to control the airflow effector device according to a nominal schedule, detect, based on a signal from the sensor indicating a vibration amplitude of the rotating blades within a frequency band, an incipient instability condition, in response to the incipient instability condition being present, determine a modified control parameter for at least one of the airflow effector device, and control the airflow effector device according to the modified control parameter, deviating from the nominal schedule.
Systems, apparatus, articles of manufacture, and methods are disclosed for a thermal management system including a first heat exchanger with a first path and a second path to transfer thermal energy between a first fluid and a second fluid; a second heat exchanger including a third path and a fourth path to transfer thermal energy between the first fluid and a third fluid; a first thermostatic element to actuate a first thermostatic valve to control a first rate of flow of the first fluid through the first heat exchanger based on a first temperature of the first fluid; and a second thermostatic element to actuate a second thermostatic valve to control a second rate of flow of the first fluid through the second heat exchanger based on a second temperature of the third fluid.
A powder handling device is provided, along with apparatus and methods of its use. The powder handling device includes a mounting portion defining a plane and comprising a mounting interface; a fixture device configured to mount the mounting portion over an opening of a powder container at a mounting interface; and a powder suction unit attached to the mounting portion and passing through the plane. The powder suction unit is movable relative to the mounting portion with at least one degree of freedom. The powder handling device may be mounted onto an opening of a powder container.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F02K 3/02 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
F04D 19/00 - Axial-flow pumps specially adapted for elastic fluids
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
B64D 35/04 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
General Electric Company Polska sp. z o.o (Poland)
General Electric Deutschland Holding GmbH (Germany)
Inventor
Schuler, Pascal
Pasieczny, Aleksander
Pritchard, Byron A.
Califf, Charles D.
Abstract
A turbine engine includes an engine plenum disposed within the turbine engine and defined at least partially by a plenum end wall and a turbo-engine disposed within the engine plenum. The turbine engine causes core air to flow through the engine plenum and into the turbo-engine. A particle deflector assembly is disposed at the plenum end wall. The particle deflector assembly includes one or more particle deflector walls that extend into the engine plenum and capture particles within the core air.
F02C 7/052 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with dust-separation devices
B01D 45/08 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by utilising inertia by impingement against baffle separators
A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.
Airfoils for turbo engines with internal vibrational dampers are disclosed herein. An example airfoil includes an outer shell body defining an interior cavity. The airfoil includes a recess formed on a surface of a structure in the interior cavity and an opening in the structure at a center of the recess. The recess results in a thickness of the structure that decreases toward the center according to a power-law profile to concentrate vibrational waves toward the center and reduce vibrations in the airfoil.
There are provided systems and methods for inversion control of turbine engines. For example, there is provided a processor-implemented method that includes a processor and a memory. The memory includes instructions which, when executed by the processor, cause the system at least to perform: simulating, by a simulated state space model, a desired dynamic response based on the sensor data and a control input; inverting the desired dynamic response as output by the state space model; determining an error between a perceived dynamic response and the inverted desired dynamic response; correcting the state space model for the determined error based on updating the one or more model parameters using a machine learning network; generating the desired dynamics based on the updated state space model; and controlling the engine based on the state space model.
G06F 30/27 - Design optimisation, verification or simulation using machine learning, e.g. artificial intelligence, neural networks, support vector machines [SVM] or training a model
A gas turbine engine is provided. The gas turbine engine includes: a fan blade and a trunnion. The trunnion defines a longitudinal direction and a transverse direction. The fan blade is coupled to the trunnion. The gas turbine engine further includes a disk. The trunnion is coupled to the disk and defines a load path from the trunnion to the disk. The gas turbine engine further includes a crushable bearing ring positioned between the trunnion and the disk along the defined load path.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
Example pump systems having dual-function annular heat exchangers are disclosed. An example pump system to pressurize a fluid within a closed loop transport bus includes a pump to move the fluid, a conduit in fluid connection with the pump, a heat exchanger positioned around at least a portion of the conduit, the heat exchanger to receive a first electrical signal transmitted in a first direction at a first time and a second electrical signal transmitted in a second direction at a second time different from the first time, the second direction opposite the first direction.
H10N 10/13 - Thermoelectric devices comprising a junction of dissimilar materials, i.e. devices exhibiting Seebeck or Peltier effects operating with only the Peltier or Seebeck effects characterised by the heat-exchanging means at the junction
99.
COMBUSTOR SIZE RATING FOR A GAS TURBINE ENGINE USING HYDROGEN FUEL
A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
A turbine engine includes a plurality of fan blades configured to rotate about a longitudinal centerline axis of the turbine engine, a pitch actuator configured to control a pitch angle of each of the plurality of fan blades, a sensor assembly configured to detect a holding torque of the plurality of fan blades, and a controller configured to monitor propeller whirl in the plurality of fan blades by detecting the holding torque. The sensor assembly includes a first sensor coupled to the pitch actuator and configured to detect the holding torque of all of the plurality of fan blades and a second sensor coupled to the pitch actuator and configured to detect the holding torque of at least one fan blade of the plurality of fan blades. The pitch actuator is configured to take a corrective action when propeller whirl is detected outside of a predetermined limit.
