A gas turbine includes a compressor section, a combustion section, and a turbine section. The combustion section includes a combustor liner at least partially defining a plurality of combustor cups including a first set of combustor cups and a second set of combustor cups. A first fuel circuit includes a first manifold fluidly coupled with a fuel source and the first set of combustor cups. A second fuel circuit includes a second manifold fluidly coupled to the second set of combustor cups. A controller is configured to detect an engine condition, and to control the first fuel circuit and the second fuel circuit to selectively provide fuel from the fuel source to at least one of the first set of combustor cups or the second set of combustor cups according to the engine condition to facilitate lightoff or relighting of the combustion section.
A coating system configured to be applied to a thermal barrier coating of an article includes an infiltration coating configured to be applied to the thermal barrier coating. The infiltration coating infiltrates at least some pores of the thermal barrier coating. The infiltration coating decomposes within at least some pores of the thermal barrier coating to coat a portion of the at least some pores of the thermal barrier coating. The infiltration coating reduces a porosity of the thermal barrier coating. The coating system also includes a reactive phase spray formulation coat configured to be applied to the thermal barrier coating. The reactive phase spray formulation coating reacts with dust deposits on the thermal barrier coating
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coatingContact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
3.
METHODS FOR CASTING A COMPONENT VIA A UNITARY CORE-SHELL MOLD
A method is provided for casting a component. Accordingly, data indicative of at least one location of a unitary core-shell mold which is susceptible to a stress concentration is received. An additive manufacturing process is employed to form the unitary core-shell mold defining a casting cavity. The unitary core-shell mold includes a shell wall defining an outer component shape and a core wall positioned inward of the shell wall. The core wall defines an inner component shape. The core wall and/or the shell wall defines at least one reinforcement recess adjacent to the at least one location which is susceptible to the stress concentration. Following the forming of the unitary core-shell mold, at least one support member is positioned within the reinforcement recess in contact with the at least one location. With the support member in place, the component is cast within the casting cavity.
A method of operating an engine is provided. The method includes operating the engine at a rated power level. Operating the engine at the rated power level includes operating the engine with a power performance indicator quantity (PIQPower) in a range of 1400 pounds per square inch (lbs/in2) to 2500 lbs/in2, wherein the power performance indicator quantity is determined according to:
A method of operating an engine is provided. The method includes operating the engine at a rated power level. Operating the engine at the rated power level includes operating the engine with a power performance indicator quantity (PIQPower) in a range of 1400 pounds per square inch (lbs/in2) to 2500 lbs/in2, wherein the power performance indicator quantity is determined according to:
PIQ
Power
=
PIQ
HPTI
PIQ
HPCE
×
N
1
N
2
×
F
n
Total
A
HPCE
.
An apparatus for additively manufacturing three-dimensional objects includes one or more laser beam sources configured to generate a laser beam, and one or more modulation devices disposed downstream of the one or more laser beam sources and configured to receive the laser beam and modulate the laser beam to generate a modulated beam. One or more computing systems are configured to generate one or more output images based on an input to the one or more computing systems of a wavefront pattern, wherein the one or more output images are generated by at least one of a graphics processing unit (GPU) or an integrated circuit. The one or more computing systems are configured to control the one or more modulation devices to generate the modulated beam with the wavefront pattern based on at least one output image of the one or more output images.
B29C 64/393 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
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
B29C 64/268 - Arrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB]
A combustion section of a turbine engine having a primary combustion chamber and a secondary combustion chamber. The primary combustion chamber is defined, at least in part, by a combustor liner and a dome wall. A set of primary fuel nozzles have outlets located at the dome wall, wherein the outlets of the set of primary fuel nozzles are fluidly coupled with the primary combustion chamber. At least one opening located in an outer liner axially aft of the dome wall fluidly couples the secondary combustion chamber and the primary combustion chamber. A set of secondary fuel nozzles are fluidly coupled with the secondary combustion chamber.
A high-speed, air-breathing propulsion engine includes a rotating detonation combustor configured to burn a fuel-air mixture and at least one ram scoop. The at least one ram scoop is configured to separate incoming air into a core flow stream and a pilot flow stream. The at least one ramp scoop is configured to cause the separation of the core flow stream and the pilot flow stream at a location disposed within a range of locations. The engine is configured to combine the pilot flow stream and the core flow stream in the rotating detonation combustor to enable a rotating detonation wave in the rotating detonation combustor.
F02K 7/14 - Plants in which the working-fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fanControl thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines with external combustion, e.g. scram-jet engines
Flowpath assemblies, methods of forming flowpath assemblies, and hypersonic vehicles are provided. For example, a flowpath assembly for a combustor comprises a tube array comprising a plurality of tubes, a joining material disposed between adjacent tubes of the plurality of tubes to join together the adjacent tubes, a flowpath layer, and an outer layer. The plurality of tubes and the joining material are disposed between the flowpath layer and the outer layer. The flowpath layer defines a combustion flowpath. Each of the plurality of tubes, the joining material, the flowpath layer, and the outer layer are formed from a composite material. The combustor comprising the flowpath assembly may be included in a ramjet engine of a hypersonic vehicle. A fabrication method may include laying up composite plies to form a tube array including the plurality of tubes, the joining material, and the flowpath and outer layers.
F02K 7/10 - Plants in which the working-fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fanControl thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or coreShaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
Methods and apparatus to control a surface of an aircraft engine are disclosed. An example system to control a surface in an aircraft engine comprises a first valve to vary a flow of cold fluid from a thermal transfer bus (TTB) to an active surface control (ASC) system based on an operating condition of the aircraft engine, the ASC system positioned adjacent to the surface, the first valve positioned upstream from the surface, and a second valve to vary a flow of hot fluid from the TTB to the ASC system based on the operating condition, the second valve positioned downstream from the surface.
A bleed valve system for a gas turbine engine includes a bleed valve, a bleed valve exhaust duct fluidly connected to the bleed valve, and a bleed valve exhaust nozzle fluidly connected to the bleed valve exhaust duct, wherein the bleed valve exhaust duct and the bleed valve exhaust nozzle are fluidly separated from an exhaust nozzle of the gas turbine engine when the bleed valve system is installed in the gas turbine engine.
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
11.
GAS TURBINE ENGINES INCLUDING EMBEDDED ELECTRICAL MACHINES AND ASSOCIATED COOLING SYSTEMS
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Pazinski, Adam Tomasz
Sobaniec, Miroslaw
Delametter, Christopher N.
Abstract
A gas turbine engine includes a fan located at a forward portion of the gas turbine engine. A compressor section and a turbine section are arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with at least a portion of the compressor section and with at least a portion of the turbine section. An electrical machine is coupled to the rotary member and is located at least partially inward of the core airflow path in a radial direction. An enclosure at least partially encloses the electrical machine. The enclosure at least partially defines a first cooling airflow path within the enclosure that at least partially defines a first cooling airflow buffer cavity at least partially around the electrical machine. The first cooling airflow path is in communication with a second cooling airflow path located outside the enclosure that at least partially defines a second cooling airflow buffer cavity at least partially around the enclosure. A cooling duct provides pressurized air to the first cooling airflow path such that the air flows along both the first cooling airflow path and the second cooling airflow path providing the first cooling airflow buffer cavity and the second cooling airflow buffer cavity.
Industrial machinery, namely industrial motors, generators and distribution equipment with integrated power electronics; power system that embeds power electronics into industrial motors, generators and distribution equipment
13.
METHODS AND APPARATUS TO PRODUCE HYDROGEN GAS TURBINE PROPULSION
Methods and apparatus to produce hydrogen gas turbine propulsion are disclosed. An example method includes activating at least one heat exchanger operatively coupled to a fuel line, injecting hydrogen into the fuel line to distribute heat from the at least one heat exchanger in the fuel line, and in response to a temperature of the fuel line being greater than a liquification temperature of an inert gas: injecting the inert gas into the fuel line, terminating injecting the hydrogen into the fuel line, and capturing the inert gas in the fuel line.
F02C 7/224 - Heating fuel before feeding to the burner
B64D 37/32 - Safety measures not otherwise provided for, e.g. preventing explosive conditions
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
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
14.
PREFORM FOR A COMPOSITE AIRFOIL OF A TURBINE ENGINE
A preform for a composite airfoil of a gas turbine engine includes a first woven fabric and a second woven fabric. The second woven fabric is located opposite the first woven fabric, with a second inner surface of the second woven fabric opposing a first inner surface of the first woven fabric to form a preform gap therebetween. A first transverse woven fabric portion extends from the first inner surface towards the second inner surface, and a second transverse woven fabric portion extends from the second inner surface towards the first inner surface. The second transverse woven fabric portion is engaged with the first transverse woven fabric portion to form a joint. Each of the first woven fabric, the second woven fabric, the first transverse woven fabric portion, and the second transverse woven fabric portion is a three-dimensional woven fabric including a plurality of reinforcing fiber tows.
A gas turbine engine including a frame, a plurality of fan blades configured to rotate about a longitudinal centerline axis of the gas turbine engine, and a mounting assembly coupling the frame to a structure. The gas turbine engine includes a sensor assembly coupled to the mounting assembly and having at least one sensor configured to detect a loading on the mounting assembly and to take a corrective action when propeller whirl is detected above a predetermined limit, and a feature configured to take a corrective action when propeller whirl is detected above a predetermined limit.
F04D 27/00 - Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
B64D 27/12 - Aircraft characterised by the type or position of power plants of gas-turbine type within, or attached to, wings
B64D 27/40 - Arrangements for mounting power plants in aircraft
F01D 7/00 - Rotors with blades adjustable in operationControl thereof
F01D 17/04 - Arrangement of sensing elements responsive to load
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for
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
F01D 21/14 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to other specific conditions
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
A turbine engine operable in a cold start condition to prevent wear on components of a gearbox assembly of the turbine engine. The turbine engine includes a gearbox assembly, a pump for directing lubricant to the gearbox assembly, a supply line heating path comprising a heat exchanger, a recirculation bypass path, a valve being positionable between a first position to direct the flow of the lubricant into the supply line heating path and a second position to direct the flow of the lubricant into the recirculation bypass path, and an electronic control unit configured to position the valve between the first position and the second position based on a temperature of the lubricant.
A gas turbine engine includes a gear assembly including planet gears arranged in a planetary configuration. Each planet gear includes a pin, a gearbox bearing, and a planet gear rim. An inner surface of the planet gear rim and an outer surface of the pin define a clearance therebetween that is greater than zero when radial, pinch, tangential, and centrifugal component forces are applied to the planet gear. Each planet gear includes a pin clearance parameter greater than or equal to zero rpm and less than or equal to 3,334 rpm. The gas turbine engine includes a primary lubrication system that supplies lubricant to the gearbox bearing of at least one planet gear during normal operation of the gas turbine engine.
A turbine engine including a fan having a plurality of fan blades, a nacelle that extends circumferentially about the fan, an engine intake including an engine inlet, and a variable engine intake system. The nacelle includes a fan cowl and an inlet cowl that is movable with respect to the fan cowl. The engine inlet is defined from a leading edge of the inlet cowl to the plurality of fan blades. The engine inlet has an inlet tip diameter at the leading edge of the inlet cowl. The variable engine intake system adjusts the inlet cowl to adjust the inlet tip diameter of the engine inlet.
An assembly for an aircraft includes a gas turbine engine comprising a turbine and a casing, and a system comprising a processor and a memory, the memory storing instructions executable by the processor to determine an exhaust gas temperature based on data from one or more sensors disposed in the gas turbine engine, the exhaust gas temperature being a temperature of an airflow downstream of the turbine, determine whether the exhaust gas temperature of the gas turbine engine is within a threshold of a specified exhaust gas temperature, and upon determining that the exhaust gas temperature is outside the threshold of the specified exhaust gas temperature, control at least one of an electric machine operably connected to the turbine or an active clearance control valve to adjust a clearance between two components of the gas turbine engine until a current exhaust gas temperature is within the threshold of the specific exhaust gas temperature.
The present invention provides a radiopharmaceutical composition comprising the following four components: (i) a radio-labelled compound; (ii) ethanol; (iii) a stabilizer of the radio-labelled compound; and (iv) a cyclodextrin.
The present invention provides a radiopharmaceutical composition comprising the following four components: (i) a radio-labelled compound; (ii) ethanol; (iii) a stabilizer of the radio-labelled compound; and (iv) a cyclodextrin.
The present invention also provides a radiopharmaceutical composition comprising: (i) a radio-labelled compound; (ii) a stabilizer of the radio-labelled compound, wherein the stabilizer comprises: ascorbic acid, aspartic acid, cysteine, maleic acid, gentisic acid, glutathione, glutamic acid, mannitol, nicotinamide, calcium chloride, N-t-butyl-alpha-phenylnitrone (PBN), tartaric acid, para-aminobenzoic acid (pABA), chloride ions or salts or combinations thereof; and (iii) a cyclodextrin.
An electrical machine comprising a rotatable shaft; an induction generator mechanically coupled to the rotatable shaft and defining a power output connectable with an electrical load, wherein the power output defines a desired constant voltage output; a converter electrically connected with the power output; and a controller connected to the converter, the controller configured to at least one of provide supplemental power at the power output or absorb excess power at the power output.
Methods and apparatus for anti-ice heat supply from waste heat recovery systems are disclosed. An apparatus for an aircraft, the apparatus comprising a fuel heat exchange system, an anti-ice heat exchanger, a waste heat recovery heat exchanger, and a conduit coupled to the fuel heat exchange system, the anti-ice heat exchanger, and the waste heat recovery heat exchanger, the conduit including a first portion and a second portion distinct from the first portion, wherein the first portion of the conduit carries a first portion of a thermal transfer fluid from the waste heat recovery heat exchanger to the anti-ice heat exchanger in which the thermal transfer fluid supplies anti-ice heat to a portion of the aircraft, wherein the second portion of the conduit carries a second portion of the thermal transfer fluid from the waste heat recovery heat exchanger to the fuel heat exchange system.
A method of manufacturing a composite component having an outer shell, an inner hub, and a plurality of struts connecting the outer shell and the inner hub. A plurality of outer shell preform portions having bifurcated strut portions are connected together to form an outer shell hoop preform, and a plurality of inner hub preform portions having bifurcated strut portions are connected together to form an inner hub hoop preform. The bifurcated strut portions are arranged to extend between the outer shell hoop preform and the inner hub hoop preform, and are arranged adjacent to one another to form a strut preform. A matrix material is injected into the mold tooling structure and a curing process is applied to the mold tooling structure to obtain the composite component.
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/46 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
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
24.
METHOD OF MANUFACTURING A COMPOSITE COMPONENT FOR A GAS TURBINE ENGINE
A method of manufacturing a composite component having an outer shell, an inner hub, and a plurality of struts connecting the outer shell and the inner hub. An outer shell frame preform is woven to include a plurality of outer shell strut preform portions, each including an outer shell strut leading edge preform portion and an outer shell strut trailing edge preform portion. An inner hub frame preform is woven to include a plurality of inner hub strut preform portions, each including an inner hub strut leading edge preform portion and an inner hub strut trailing edge preform portion. In forming a strut preform, the outer shell leading edge preform portion and the inner hub leading edge preform portion are connected together, and the outer shell trailing edge preform portion and the inner hub trailing edge preform portion are connected together.
A frame assembly includes a frame and a mounting bracket assembly. The frame includes an inner hub, an outer shell located opposite the inner hub, and a plurality of struts connecting the inner hub with the outer shell. At least one strut of the plurality of struts is a hollow strut with a radial passage extending therethrough. The mounting bracket assembly includes an inner flange located on an inner surface of the inner hub, an outer flange located on an outer surface of the outer shell, and a radial linkage connecting the inner flange with the outer flange. The radial linkage extends through the radial passage of the hollow strut.
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 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
27.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
GB
.
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
GB
.
VG is a gutter volume of the gutter and VGB is a gearbox volume.
An aircraft including an engine defining a centerline axis, the engine comprising a fan and a turbomachine rotatably driving the fan, the turbomachine including an exhaust section comprising an outlet nozzle, the outlet nozzle including a fixed portion and a movable portion, wherein the movable portion is movable from a first position to a second position, wherein when the movable portion is in the first position, the movable portion is aligned with the centerline axis, and wherein when the movable portion is in the second position, the movable portion is canted downward in the vertical direction and outward in the lateral direction relative to the centerline axis, and wherein the movable portion includes an end engaging the fixed portion, wherein when the movable portion is in the first position, the end defines a nonzero cant angle with the centerline axis.
Trunnions having multiple fan blades for use with an aircraft engine are disclosed herein. An example comprises a trunnion rotatable within a hub of an aircraft engine, a first fan blade coupled to the trunnion, the first fan blade having a pitch axis, and a second fan blade coupled to the trunnion, the second fan blade having a longitudinal axis, the longitudinal axis laterally offset from the pitch axis.
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
F04D 19/00 - Axial-flow pumps specially adapted for elastic fluids
Methods and apparatus to reduce vibration of a radial seal in a gas turbine engine are disclosed. An example apparatus includes a seal segment of a radial seal positioned between a rotor and a stator of a gas turbine engine, the seal segment including a first seal body adjacent the rotor and a first surface of the stator, a second seal body adjacent a second surface of the stator, the second surface opposite the first surface, the second seal body pivotable relative to the first seal body, and a pivot bar extending between the first seal body and the second seal body, and a damper operatively coupled to the pivot bar.
F01D 11/08 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator
F03G 7/06 - Mechanical-power-producing mechanisms, not otherwise provided for or using energy sources not otherwise provided for using expansion or contraction of bodies due to heating, cooling, moistening, drying, or the like
31.
ADDITIVE MANUFACTURING APPARATUSES WITH REMOVABLE BUILD BOXES AND BUILD BOX MANAGEMENT SYSTEMS
An apparatus for forming a three-dimensional article having successive layers of a powder material, which layers correspond to successive cross-sections of the three-dimensional article includes a process chamber comprising a process chamber floor and a build box management system. The build box management system includes an intake conveyor that receives a build box and an elevated conveyor that engages the build box at a location above a bottom of the build box.
B29C 64/259 - Enclosures for the building material, e.g. powder containers interchangeable
B29C 64/165 - Processes of additive manufacturing using a combination of solid and fluid materials, e.g. a powder selectively bound by a liquid binder, catalyst, inhibitor or energy absorber
B29C 64/307 - Handling of material to be used in additive manufacturing
Methods and apparatus for sensor-based part development are disclosed. An example apparatus includes at least one memory, instructions in the apparatus, and processor circuitry to execute the instructions to translate at least one user-defined material property selection into a desired process observable, the desired process observable including a meltpool property, perform voxel-based autozoning of an input part geometry, the input part geometry based on a computer-generated design, and output a voxelized reference map for the input part geometry based on the desired process observable and the voxel-based autozoning.
A system for preventing asynchronous rotation in aircraft components includes an optical emitting source, an optical receiver, and a retroreflector configured to be disposed on a rotating component connected to an engine of an aircraft. The system emits radiant flux from the optical emitting source towards the retroreflector when the retroreflector is on the rotating component and the engine positions the rotating component into a field of view of the optical emitting source; receives incident radiant flux from the retroreflector by the optical receiver when the retroreflector is within the field of view of the optical emitting source; records a time of receiving the incident radiant flux; determines a rotational speed of the rotating component based on the recorded time; determines whether the rotational speed is within a predetermined range; and selectively changes a state of the engine when the rotational speed is outside of the predetermined range.
A method for additively manufacturing three-dimensional objects includes generating a laser beam with one or more laser beam sources and modulating the laser beam via one or more modulation devices to generate a modulated beam based on a set of one or more defined parameters where the modulated beam is directed onto a build plane. One or more sensor devices detect first data corresponding to one or more measured parameters of the modulated beam and second data corresponding to reflected radiation from the build plane. The set of the one or more defined parameters is modified based on the first data and the second data.
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/36 - Process control of energy beam parameters
B22F 10/85 - Data acquisition or data processing for controlling or regulating additive manufacturing processes
B22F 12/43 - Radiation means characterised by the type, e.g. laser or electron beam pulsedRadiation means characterised by the type, e.g. laser or electron beam frequency modulated
B22F 12/90 - Means for process control, e.g. cameras or sensors
B28B 1/00 - Producing shaped articles from the material
B28B 17/00 - Details of, or accessories for, apparatus for shaping the materialAuxiliary measures taken in connection with such shaping
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
B29C 64/273 - Arrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] pulsedArrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] frequency modulated
B29C 64/393 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
A control circuit determines a first period of time during which a jet turbine engine-powered aircraft was operating, then aggregates data from a plurality of data sources including an Aircraft Condition Monitoring System and an Electronic Engine Control that corresponds to that first period of time to provide aggregated data, and then determines whether the aggregated data meets a standard for quantitative sufficiency. By one approach, determining whether the aggregated data meets a standard for quantitative sufficiency can comprise determining whether the aggregated data meets a standard for quantitative sufficiency as a function of assessing the intermittency of data from at least some of the plurality of data sources. Assessing that intermittency of data from at least some of the plurality of data sources can comprise comparing received data to corresponding intermittency signatures.
A gas turbine engine includes a core engine coupled to an input shaft, a fan coupled to an output shaft, and a gearbox assembly. The gearbox assembly includes a split sun gear coupled to the input shaft, a plurality of planet gears intermeshing with the split sun gear, and a ring gear intermeshing with the planet gears. The split sun gear includes a forward sun gear and an aft sun gear each coupled to the input shaft. Each planet gear includes a pin with an outer surface, a rim with an inner surface, and a clearance between the pin's outer surface and the rim's outer surface. The clearance is greater than zero when radial, pinch, tangential, and centrifugal component forces are applied to the planet gear. Each planet gear includes a pin clearance parameter greater than or equal to zero rpm and less than or equal to 3,334 rpm.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
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
37.
TURBINE ENGINE WITH COMPOSITE AIRFOIL HAVING A NON-METALLIC LEADING EDGE PROTECTIVE WRAP
A composite airfoil having a non-metallic leading edge protective wrap is provided. In one aspect, the airfoil has a composite core having a pressure sidewall and a suction sidewall each extending between a core leading edge and a core trailing edge. A leading edge protective wrap protects the core leading edge and includes a trailing wrap and a leading wrap. The trailing wrap wraps around the core leading edge and is connected to the composite core. The leading wrap wraps around the core leading edge and is connected to the trailing wrap. The trailing and leading wraps both have leading edges that are spaced from one another. A filler is positioned between the leading edges of the trailing and leading wraps. A protective nose is connected to the leading edge of the leading wrap. The components of the leading edge protective wrap are formed of non-metallic materials.
A gas turbine engine includes a fan, a core engine coupled to the fan, a fan case housing the fan and the core engine, a plurality of outlet guide vanes extending between the core engine and the fan case, and an acoustic spacing. The fan includes a plurality of fan blades that define a fan diameter and a BEAL. The acoustic spacing includes a distance between the fan and the outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The combination of the acoustic spacing and a corrected specific thrust of the gas turbine engine provide enhanced propulsive efficiency.
A power transfer system includes a first pressure spool of a gas turbine engine structured to rotate and compress a working fluid to a first pressure, a first electric machine connected to the first pressure spool, the first electric machine including a first stator, a second pressure spool of the gas turbine engine structured to rotate and compress the working fluid to a second pressure different than the first pressure, a second electric machine connected to the second pressure spool, the second electric machine including a second stator, a first power converter connected to one of the first electric machine or the second electric machine, and an electrical connector including a main portion connecting the first stator of the first electric machine to the second stator of the second electric machine while bypassing the first power converter and a branch portion connecting the main portion to the first power converter.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
H02K 7/18 - Structural association of electric generators with mechanical driving motors, e.g.with turbines
H02K 7/20 - Structural association with auxiliary dynamo-electric machines, e.g. with electric starter motors or exciters
H02K 11/04 - Structural association of dynamo-electric machines with electric components or with devices for shielding, monitoring or protection for rectification
A lubrication system for a turbine engine. The turbine engine includes a fan having a fan shaft and one or more rotating components. The lubrication system includes one or more tanks that stores lubricant therein, 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 normal operation of the turbine engine. The auxiliary lubrication system includes an auxiliary pump that is coupled to the fan shaft. The auxiliary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components based on a pressure of the lubricant in the primary lubrication system. Rotation of the fan shaft causes the auxiliary pump to pump the lubricant to the one or more rotating components.
A system including a test component including a first gas channel formed in the test component, a surface at least partially coated with pressure sensitive paint, and seed channels formed in the test component extending from the first gas channel to the surface, the seed channels defining seed holes formed in the surface opposite the first gas channel. A first gas source delivers a first gas to the first gas channel, a second gas source delivers a second gas across the seed holes, an imaging device capturing image data of a seed flow streak of the first gas based on a change in luminescence intensity of the pressure sensitive paint, and an electronic control unit determines whether an angle between the seed flow streak and an imaginary line extending parallel to a centerline of the test component is greater than a predetermined threshold.
An apparatus for additively manufacturing three-dimensional objects includes a laser beam source configured to generate a laser beam, and a modulation device disposed downstream from the laser beam source. The modulation device includes a dichroic layer and a modulation layer. The dichroic layer is configured to transmit a first wavelength range of the laser beam and a second wavelength range of the laser beam to the modulation layer and reject wavelengths of the laser beam falling outside the first and second wavelength ranges where the first wavelength range is a multiple of the second wavelength range. One or more optical devices are disposed downstream of the modulation device and configured to selectively scan the laser beam onto a build plane.
B29C 64/273 - Arrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] pulsedArrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] frequency modulated
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/36 - Process control of energy beam parameters
B22F 12/43 - Radiation means characterised by the type, e.g. laser or electron beam pulsedRadiation means characterised by the type, e.g. laser or electron beam frequency modulated
B28B 1/00 - Producing shaped articles from the material
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
A method for additively manufacturing three-dimensional objects includes generating a laser beam with a laser beam source. A primary laser beam emitted along a beam path is generated using a first polarization direction of the laser beam unmodulated by a modulation device positioned downstream of the laser beam source, and a secondary laser beam emitted along the beam path is generated by modulating a second polarization direction of the laser beam via the modulation device. The primary laser beam and the secondary laser beam are coupled and directed, via a deflection device positioned downstream of the modulation device, onto a powder build material supported by a build platform.
G02F 1/01 - Devices or arrangements for the control of the intensity, colour, phase, polarisation or direction of light arriving from an independent light source, e.g. switching, gating or modulatingNon-linear optics for the control of the intensity, phase, polarisation or colour
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/362 - Process control of energy beam parameters for preheating
B22F 12/41 - Radiation means characterised by the type, e.g. laser or electron beam
B22F 12/90 - Means for process control, e.g. cameras or sensors
B28B 1/00 - Producing shaped articles from the material
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
B29C 64/273 - Arrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] pulsedArrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB] frequency modulated
Methods and apparatus to vary geometry of a thrust link are disclosed. An apparatus to support an aircraft engine includes a thrust link including a forward end coupled to the aircraft engine, an aft end coupled to the aircraft engine or a pylon, and a bumper coupled to (i) the thrust link between the forward end and the aft end and (ii) to an annular fan casing, a nacelle, the pylon, or an aircraft associated with the aircraft engine, a bumper distance percentage defined between the aft end and a location on the thrust link at which the bumper couples to the thrust link, wherein the location is aft of the forward end.
A flexible pipe for conveying liquid hydrogen. The flexible pipe includes a multilayer wall having an inner layer, an outer layer, and a plurality of insulating layers between the inner layer and the outer layer. The inner layer is a hydrogen barrier layer and defines a flow passage for conveying liquid hydrogen. The outer layer forms an exterior of the pipe. The outer layer includes a plurality of reinforcing fiber tows.
F16L 11/08 - Hoses, i.e. flexible pipes made of rubber or flexible plastics with reinforcements embedded in the wall
G01M 3/18 - Investigating fluid tightness of structures by using fluid or vacuum by detecting the presence of fluid at the leakage point using electric detection means for pipes, cables, or tubesInvestigating fluid tightness of structures by using fluid or vacuum by detecting the presence of fluid at the leakage point using electric detection means for pipe joints or sealsInvestigating fluid tightness of structures by using fluid or vacuum by detecting the presence of fluid at the leakage point using electric detection means for valves
A composite hydrogen storage tank for storing liquid hydrogen incudes a multilayer composite wall. The multilayer composite wall defines a chamber and includes an inner composite layer, an outer composite layer, and a plurality of insulating layers between the inner composite layer and the outer composite layer. A hydrogen barrier layer is formed on an inner surface of the inner composite layer. Each of the inner composite layer and the outer composite layer including a plurality of reinforcing fiber tows surrounded by a matrix. The outer composite layer forms an exterior of the composite hydrogen storage tank.
A composite storage tank for liquid hydrogen and method of manufacturing the composite storage tank. The composite storage tank for liquid hydrogen includes a vessel wall defining a chamber to hold liquid hydrogen. A buffer can be located in the chamber. The buffer includes a plurality of microchannel flow passages that fluidly connect a first side with a second side. The buffer can be connected to the vessel wall by a receiver integrally formed in the vessel wall. The composite storage tank can include a vessel wall and a receiver integrally formed in the vessel wall. The method can include laying up a plurality of reinforcing fiber tows on a layup tool to integrate a flange portion of the receiver with the plurality of reinforcing fiber tows. The method can also include inserting a stiffener into a gap of the receiver. The stiffener can be the buffer.
A low pitch protection control system and method that includes sensing hardware, an engine controller, and a pitch actuation system for variable pitch fan blades of a turbine engine is provided. The engine controller receives sensed data from the sensing hardware related to current flight conditions, determines, based on at least some of the sensed data, a first pitch angle threshold value for initiating a first mitigating action to prevent onset of excessive drag at the current conditions, monitors a current pitch angle of one or more variable pitch fan blades of the turbine engine relative to the first pitch angle threshold value, and commands the pitch actuation system to perform the first mitigating action in response to the current pitch angle being less than or equal to the first pitch angle value.
A fan rotor bearing support system for a turbine engine having a power gearbox includes a first bearing assembly mounted on a proximal portion of a fan rotor of a fan section of the turbine engine and a second bearing assembly mounted on a more distal portion of the fan rotor proximal of a fan hub of the fan section. The first bearing assembly includes a four point ball bearing with limited clearance between the balls of the bearing assembly and the races of the bearing assembly to limit axial movement of the fan rotor in relation to the power gearbox. The second bearing assembly includes a low profile roller bearing to facilitate a reduction in the fan hub ratio of the turbine engine.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F16C 19/16 - Bearings with rolling contact, for exclusively rotary movement with bearing balls essentially of the same size in one or more circular rows for both radial and axial load with a single row of balls
A combustion section for a turbine engine. The combustion section has a circumferential casing, a combustor, multiple fuel nozzles and a fuel supply system. The circumferential casing defines an interior. The combustion has an annular liner located within the interior. The fuel supply system has a fuel manifold and multiple fuel line branches.
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.
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.
A system for using large language models (LLMs) to convert natural language (NL) requirements for aerospace systems to formalism for controlling a landing gear system of an aircraft includes a processor and a memory including instructions which, when executed by the processor, causes the system at least to perform using an LLM to revise a received NL expression associated with a requirement for an aerospace system. An LLM is also used to replace at least one word in the revised NL expression with a lifted atomic entity to generate an underlying sentence structure and translate the underlying sentence structure to a lifted logical formula. An LLM is also used to ground the lifted logical formula and to convert the NL expression into a formalized requirement for the aerospace system. The system also controls the aerospace system based on the formalized requirement.
A method for processing a metal article includes pressure heating the metal article at a first temperature that is below a solvus temperature of a precipitate to reduce growth of surface variation of an outer surface of the metal article and heating the metal article at a second temperature that is above the solvus temperature of the precipitate to form a recrystallized grain structure of the metal article.
C22F 1/10 - Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of nickel or cobalt or alloys based thereon
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B33Y 80/00 - Products made by additive manufacturing
C22C 19/05 - Alloys based on nickel or cobalt based on nickel with chromium
A gearbox assembly includes a planet gear including an outer surface and an inner surface opposite the outer surface. A pair of circumferential oil channels are formed in the inner surface proximate opposite ends of the planet gear. A distributor channel is formed to extend in a longitudinal direction toward a center of the planet gear from each circumferential oil channel. One or more tooth channels are formed to extend from each distributor channel to the outer surface of the planet gear. A journal pin is received within the planet gear extending along an axis about which the planet gear rotates. A gap defines a journal bearing formed between the inner surface of the planet gear and the journal pin. An oil passage is provided supplying oil to the gap.
A gas turbine engine defines a radial direction and an axial direction. The gas turbine engine includes an inlet duct comprising a splitter, a fan duct downstream from the inlet duct and the splitter, and a core duct downstream of the inlet duct and the splitter. The core duct may be radially inward of the fan duct. A heat exchanger assembly may extend annularly about at least one of the inlet duct, the fan duct, or the core duct. The heat exchanger assembly may include a sheet extending circumferentially about the at least one of the inlet duct, the fan duct, and the core duct. One or more aerodynamic features extend from the sheet, the one or more aerodynamic features within the at least one of the inlet duct, the fan duct, and the core duct.
A power switching system can include a main power supply to provide power to electrical devices, such as electric heaters for ice protection during an aircraft operation. During a nominal mode of operation, a subset of electrical devices may be operated sufficiently based on the main power supply. The power switching system may further include an auxiliary power supply to supplement a main power supply for critical phases of operation, such as during takeoff, landing, or flight into known icing conditions. The auxiliary power supply can include a battery or supercapacitor that augments electrical power provided from the main power supply. When an auxiliary mode of operation is needed, such as during the critical phase of flight, an electrical power from an auxiliary power supply can be used to augment the main power supply so that all electrical devices are powered.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a first upper lobe and a first lower lobe defining a first intervening recess, and a second upper lobe and a second lower lobe defining a second intervening recess, collectively defining a neck for mounting to the disk.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a set of lobes defining complementary recesses, collectively defining a neck for mounting to the disk.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a first upper lobe and a first lower lobe defining a first intervening recess, and a second upper lobe and a second lower lobe defining a second intervening recess, collectively defining a neck for mounting to the disk.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a set of lobes defining complementary recesses, collectively defining a neck for mounting to the disk.
A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan blades feature a low aspect ratio, reducing blade count while maintaining thrust and efficiency. Efficiency is enhanced through a determined relationship between fan blade count, aspect ratio, and specific 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
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.
A turbomachine engine includes a fan section having a fan shaft, and a core engine having one or more compressor sections, one or more turbine sections that includes a power turbine, and a combustion chamber in flow communication with the compressor sections and turbine sections. The turbomachine engine includes a low-speed shaft coupled to the power turbine and having a midshaft that extends from a forward bearing to an aft bearing. The low-speed shaft is characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-speed shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine includes a gearbox assembly that couples the fan shaft to the low-speed shaft and characterized by a gearbox assembly mode less than 95% of a midshaft mode of the midshaft or greater than 105% of the midshaft mode.
Method for forming a protected component are provided. The method may include: applying a plurality of particles in a solid state using a thermal spray process onto a surface of a substrate to form an array of discontinuous environmentally reactive deposits, where the substrate comprises an aluminum-based alloy. The array of discontinuous environmentally reactive deposits may comprise at least one reactive element comprising Al, Zn, V, Ce, Dy, Er, Eu, Gd, Ho, La, Lu, Nd, Pr, Sc, Sm, Tb, Tm, Y, Yb, Mo, Si, Ca, W, a mixture thereof, or alloys thereof.
A trained machine learning model identifies that a real-world apparatus has a failed component, which trained machine learning model has been trained with a training corpus that includes content generated by synthesizing a plurality of synthesized operating examples for a given apparatus, wherein at least some of the plurality of synthesized operating examples are generated via a simulation modeling environment that receives as input characterizing information that corresponds to any of a variety of failure states for a component of the given apparatus.
A rotor assembly for an electric machine includes a first end and a second end, the second end distal from the first end in an axial direction, a first radial wall extending between the first end and the second end and defining an inner cavity, and a second radial wall extending between the first end and the second end, the second radial wall radially-overlying the first radial wall and defining an outer cavity between the first radial wall and the second radial wall.
H02K 7/18 - Structural association of electric generators with mechanical driving motors, e.g.with turbines
H02K 9/19 - Arrangements for cooling or ventilating for machines with closed casing and closed-circuit cooling using a liquid cooling medium, e.g. oil
H02K 11/04 - Structural association of dynamo-electric machines with electric components or with devices for shielding, monitoring or protection for rectification
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.
A turbine engine includes a fan including a fan rotor assembly having a plurality of fan blades coupled to a fan disk, a fan frame that supports the fan, and a fan speed sensor system. The fan speed sensor system includes one or more fan speed sensors that sense a fan speed of the fan. The one or more fan speed sensors are positioned to sense the fan rotor assembly. A method of operating the turbine engine includes rotating the fan rotor assembly, sensing rotation of the fan rotor assembly with the one or more fan speed sensors, determining the fan speed of the fan based on the rotation of the fan rotor assembly, and controlling one or more components of the turbine engine based on the fan speed of the fan.
A casting apparatus includes a heater including a heating wall defining an interior heating chamber, a material supply disposed in the interior heating chamber, the heater configured to provide heat to the material supply, a cooler disposed beneath the heater, the cooler including a cooling wall defining an interior cooling chamber and a cooling plate configured to support a plurality of molds on an upper surface thereof, a baffle disposed between the heater and the cooler, the baffle attached to one or both of the heating wall or the cooling wall and extending across the heater and cooler thereby separating the interior cooling chamber from the interior heating chamber, the baffle including a plurality of openings, each one of the plurality of openings configured to receive a respective one of the plurality of molds.
An insertion tool for inspection includes a tube with a flexible section having at least two joints for articulating the flexible section into different shapes, the flexible section selectively configurable between an unrigidized state and at least a first rigid state having a first non-linear shape. The tool includes an end effector coupled to the tube with a sensor head to direct a beam toward a target area to perform a profilometry operation. The tool includes a waveguide extending through the tube, to transmit the beam from an electromagnetic source to the end effector and transmit electromagnetic signals from the end effector to a profilometry detecting device, and an end effector actuator to move the beam at the target. The end effector is axially spaced from an insertion axis of the tool when the flexible section is in the first rigid state to obtain off-axis profilometry measurements at the target.
The present disclosure is directed towards using lyophilized reagent beads for a fully closed multi-step reaction to complete antibiotic susceptibility tests and amplification assays. The lyophilized reagent beads include a lyophilized amplification reagent bead and an lyophilized antibiotic bead. The lyophilized reagent bead may be embedded in solid wax layer inside the test tube to enable cell growth in antibiotic prior to release of amplification reagents. A microfluidic manifold device may be used to insert the sample into the amplification tube and enable a fully closed system.
C12Q 1/689 - Nucleic acid products used in the analysis of nucleic acids, e.g. primers or probes for detection or identification of organisms for bacteria
A hydrogen fuel system for an aircraft includes a housing, a hydrogen fuel component, and a hydrogen storage structure. The hydrogen fuel component has a hydrogen fuel passage for hydrogen fuel to flow therethrough. The hydrogen fuel component is located within a volume of the housing. The hydrogen storage structure includes a hydrogen storage material having a morphology for retaining hydrogen within the hydrogen storage material. A hydrogen leak path is defined within the volume of the housing, a hydrogen accumulation region is defined within the volume of the housing, or both the hydrogen leak path and the hydrogen accumulation region is defined within the volume of the housing. The hydrogen storage structure is within the volume of the housing in at least one of the hydrogen leak path or in the hydrogen accumulation region.
A gas turbine engine includes a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a compressor, the turbine section having one or more turbines. The turbomachine includes one or more additively manufactured metal components, and a total mass of the one or more additively manufactured metal components is in a range from 1% to 50% of a total dry mass of the gas turbine engine.
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
B33Y 80/00 - Products made by additive manufacturing
75.
TURBINE ENGINE HAVING A VARIABLE ENGINE INTAKE SYSTEM
A turbine engine including a fan having a plurality of fan blades, a nacelle that extends circumferentially about the fan, an engine intake including an engine inlet, and a variable engine intake system. The nacelle includes a fan cowl and an inlet cowl that is movable with respect to the fan cowl. The engine inlet is defined from a leading edge of the nacelle to the plurality of fan blades. The variable engine intake system adjusts the inlet cowl axially between a fully retracted position and a fully extended position to adjust an inlet length of the engine inlet. The inlet cowl maintains contact with the fan cowl when the inlet cowl is extended.
A coated component for coke abatement in a gas turbine engine. The coated component includes a substrate, and a catalytic coating. The catalytic coating has a first layer in contact with the substrate, the first layer including a first ceramic material and a first noble metal, a second layer in contact with the first layer, the second layer including a second ceramic material and a second noble metal, and a third layer in contact with the second layer, the third layer including a third ceramic material and a third noble metal. The first layer has a plurality of grains having an average diameter, and the first layer has a volume percent porosity. The second layer has a plurality of grains having a bimodal distribution, and the second layer has a volume percent porosity. The third layer has a plurality of grains having an average diameter, and a volume percent porosity.
A gas turbine engine, comprising: a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body defining a fuel nozzle assembly centerline, the fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.
A turbine engine can combust fuel to drive a turbine, which drives the engine. A fuel nozzle can supply fuel for combustion or ignition of the fuel. The fuel nozzle can include a central body defining a first fuel passage terminating at a first orifice. An annular first wall circumferentially surrounds the central body and is spaced therefrom to define a first airflow passage therebetween. A set of first swirler vanes is circumferentially spaced about the central body and disposed within the first airflow passage. The fuel nozzle can supply a mixture of fuel and air for combustion.
A blended wing aircraft is provided defining a longitudinal direction, a lateral direction, and a longitudinal centerline extending along the longitudinal direction, the blended wing aircraft comprising: a body defining a leading portion; 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 an icing management system comprising a distribution module and a plurality of icing management modules in thermal communication with the leading portion of the body, the leading edges of the pair of wings, or both, the distribution module in selective communication with each of the plurality of icing management modules to alternatingly activate the plurality of icing management modules.
Lift fan assembly for an aircraft, including a rotor and an inlet nozzle. The rotor includes a fan having a plurality of fan blades and a plurality of outer airfoils. The plurality of outer airfoils is connected to the fan blades to rotate the plurality of fan blades. The inlet nozzle is positioned and oriented to direct rotating air in a rotating airflow direction to impinge on outer airfoils of the plurality of outer airfoils and to rotate the plurality of outer airfoils. The rotating airflow direction can be transverse to a radial direction and have a component in a direction tangential to the circumferential direction. The fan blades can be rotatable to generate a column of propulsor discharge air, and the lift fan assembly can include a discharge assembly with a discharge nozzle positioned to discharge the rotating air into the column of propulsor discharge air.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A turbine nozzle is arranged in the turbine section. Vanes of the turbine nozzle include a vane airfoil having cooling features formed in a trailing edge of the vane airfoil. Vanes of the turbine nozzle further include a baffle insert in a cavity of the vane airfoil. The baffle insert includes cooling holes.
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.
A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.
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.
A fuel injector for a gas turbine engine includes a compression section, combustion section, and turbine section in serial flow arrangement. The fuel injector arranged in the combustion section and includes an inner nozzle defining a fuel injector axis and an outer nozzle in annular arrangement about the inner nozzle. The outer nozzle at least partially defines a mixing region. An inner air passage and an outer air passage exhaust to the mixing region. A fuel body positions between the inner air passage and the outer air passage and has an outer fuel passage exhausting to the mixing region.
A gas turbine engine, comprising: a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.
A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; and a fuel injector assembly comprising a mixing tube having a mixing tube body defining a mixing channel. The mixing tube body can have a set of fuel passages terminating in fuel orifices fluidly coupled to the mixing channel. A set air flow passages terminating in air outlets can be fluidly coupled to the mixing channel. At least some of the air outlets circumscribe a corresponding fuel orifice.
A turbine engine, comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section further comprising a combustor liner at least partially defining a combustion chamber, and a gaseous fuel nozzle assembly. The gaseous fuel nozzle assembly comprises a set of mixing tubes, a set of fuel jets and an air passage, the air passage including a set of air passage outlets, with each air passage of the set of air passage outlets tangentially, fluidly coupled to a corresponding mixing tube air inlet.
A composite airfoil assembly and method of forming includes a spar assembly with a spar core and a spar fiber layer at least partially surrounding the spar core, as well as a root assembly with a sleeve assembly carrying a set of spaced wedges defining at least one slot configured to receive a first end of the spar assembly.
A rotor system for a turbine engine. The rotor system includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that rotate. The stator assembly includes a plurality of stator vanes arranged circumferentially about the stator assembly and includes at least one pair of non-uniform gaps between adjacent stator vanes. The plurality of stator vanes includes a first group of stator vanes having a first non-uniform gap between adjacent stator vanes, a second group of stator vanes having a second non-uniform gap between adjacent stator vanes, and a third group of stator vanes having a uniform spacing between adjacent stator vanes. The first non-uniform gap is positioned 180° from the second non-uniform gap. The plurality of rotor blades directs air through the plurality of stator vanes.
A high entropy alloy-based composition is provided that has the formula: (M1aM2bM3cM4dM5eM6f)CrAlY1-x-zZrxMoz where: each of M1, M2, M3, M4, M5, and M6 is a different alloying element selected from the group consisting of Ni, Co, Fe, Si, Mn, and Cu such that none of M1, M2, M3, M4, M5, and M6 are the same alloying element; 0.05≤a≤0.35; 0.05≤b≤0.35; 0.05≤c≤0.35; 0.05≤d≤0.35; 0.05≤e≤0.35; 0≤f≤0.35; a+b+c+d+e+f=1; 0≤x≤1; 0≤z≤1; and 0≤x+z≤1.
C23C 30/00 - Coating with metallic material characterised only by the composition of the metallic material, i.e. not characterised by the coating process
92.
VISCOUS DAMPER APPARATUS AND ASSOCIATED METHODS TO CONTROL A RESPONSE TO A RESONANT VIBRATION FREQUENCY
Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency are disclosed. An apparatus to support an aircraft engine includes a first thrust link including a forward end and an aft end, the forward end of the first thrust link coupled to the aircraft engine, a second thrust link including a forward end and an aft end, the forward end of the second thrust link coupled to the aircraft engine, and a damper coupled to the aft end of the first thrust link and to the aft end of the second thrust link.
A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.
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 method of assembling a gas turbine assembly includes providing a core gas turbine engine including a high-pressure compressor, a combustor, and a turbine, coupling a. A low-pressure turbine is axially aft from the core gas turbine engine, coupling a fan assembly is axially forward from the core gas turbine engine, and coupling a booster compressor is coupled to the low-pressure turbine such that the booster compressor and the low-pressure turbine rotate at a first rotational speed, and an epicyclic gearbox is coupled to the low-pressure turbine and the fan assembly such that the fan assembly rotates at a second rotational speed.
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
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
95.
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.
An aircraft defining a vertical direction, a longitudinal direction, and a transverse direction is provided. The aircraft includes: a fuselage; a propulsion system comprising a power source and a plurality of vertical thrust electric fans driven by the power source; and a wing extending from the fuselage in the transverse direction. The wing includes a support structure that comprises a plurality of first support members and a plurality of second support members. The plurality of first support members extending at least partially between the plurality of second support members. The plurality of vertical thrust electric fans arranged between the plurality of first support members and the plurality of second support members.
An airfoil assembly has an airfoil portion and a spar. The airfoil portion has an outer wall. The outer wall extends between a root and a tip in a spanwise direction, and between a leading edge and a trailing edge in a chordwise direction. The outer wall defines an interior. The spar has a spar body. The spar is coupled to the airfoil portion. The spar extends into the interior.
A method of decentralized collaborative target searching, target tracking, or both, includes initializing a probability distribution over a probabilistic search area with an a priori probability distribution of target locations, transmitting the probability distribution to a plurality of vehicles, capturing, using a plurality of sensors on the plurality of vehicles, sensor data relating to locations of a target, updating the a priori probability distribution based on the sensor data or based on data observations outside of the plurality of vehicles, or both, to provide an updated probability distribution, determining optimal trajectories for the plurality of vehicles using the updated probability distribution and using an ergodic principle to define optimality and based on each vehicle of the plurality of vehicles calculating its own optimal trajectory to enable decentralized calculations, and detecting or tracking, or both, the target based on the determining of the optimal trajectories.
G06N 7/01 - Probabilistic graphical models, e.g. probabilistic networks
H04W 4/46 - Services specially adapted for particular environments, situations or purposes for vehicles, e.g. vehicle-to-pedestrians [V2P] for vehicle-to-vehicle communication [V2V]
Resuscitation apparatus, namely, a component of an infant warmer, namely, an advanced configuration solution set to support resuscitation efforts featuring 3-lead electrocardiograph monitoring, blood oxygen saturation monitoring, integrated resuscitation, scale and Apgar time.
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.
