A combustor comprising a dome wall, an annular liner, a combustion chamber, a set of fuel cups, and a set of dilution passages for each fuel cup of the set of fuel cups. The set of fuel cups circumferentially spaced along the dome wall relative to the combustor centerline. The set of dilution passages terminating in a plurality of slots spaced about the corresponding fuel cup in the set of fuel cups.
Systems, apparatus, articles of manufacture, and methods are disclosed for reference markers for engine components. An example turbine engine includes a substrate and a blade coupled to the substrate, the blade including a first side having an external surface and a reference marker provided on the external surface, the reference marker including spatial marking features having predetermined dimensions, the spatial marking features including (a) a first spatial marking feature at a first location on the external surface and (b) a second spatial marking feature at a second location on the external surface that is different than the first location, wherein a combination of the first spatial marking feature and the second spatial marking feature provide a first measure of the blade based on the predetermined dimensions.
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 method of imaging a turbine engine component with a thermographic sensor, the turbine engine component having a first surface and a second surface spaced from the first surface, and a plurality of holes with inlets formed in the second surface and outlets formed in the first surface. The method including flowing air through the turbine engine component and obtaining, during the flowing of air, thermographic data when the turbine engine component and a thermograph sensor are positioned such that a viewing plane and a centerline of the outlet or a centerline of a cooling hole form and angle in a range from 60° to 120°. A subset of the thermographic data is determined and a flow score for the cooling hole is calculated, wherein the thermographic data subset, the flow score, or the thermographic data subset and the flow score are visually displayed.
A method of operating a rotating detonation combustor includes providing a flow of air through an air inlet into a detonation chamber, providing fuel from a fuel injector into one of the air inlet or the detonation chamber, mixing the fuel and air to generate a fuel-air mixture, detonating the fuel-air mixture to generate rotating detonation waves, and controlling, during operation of the rotating detonation combustor from a first power operating state to a second power operating state, different from the first power operating state, a flow of an auxiliary gas through an auxiliary gas injection port into one of the air inlet to mix with the flow of the air within the air inlet to control a discharge coefficient of the air inlet, or into the detonation chamber to control a discharge coefficient of the detonation chamber, and to control an operating mode of the rotating detonation combustor.
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
6.
COMBUSTOR DOME COUPLED TO INNER AND OUTER COMBUSTOR LINERS
A combustor dome assembly includes an inner liner, an outer liner, a dome disposed on the inner liner and the outer liner, and a cowl connected to the inner liner and to the outer liner, wherein the dome is disconnected from the cowl and defines an annular surface facing the cowl, the annular surface extending from an outer diameter of the dome to an inner diameter of the dome.
A turbine engine includes a rotor, a stator having a carrier, and a seal assembly that is disposed between the rotor and the stator. The seal assembly includes a plurality of seal segments. The plurality of seal segments includes a seal segment having a seal face forming a fluid bearing with the rotor, a body, and an aft bearing extending from the body. The turbine engine further includes a roller assembly having one or more rolling elements coupled to one of the aft bearing or the carrier. The one or more rolling elements in rolling contact with the other of the aft bearing or the carrier.
A build material escapement assembly for an additive manufacturing system includes a base defining a cavity and an aperture opening into the cavity, a diaphragm extending at least partially across the aperture and coupled to the base, and a plate disposed within the cavity and coupled to the diaphragm. The build material escapement assembly further includes a post disposed within the cavity beneath the diaphragm and coupled to the plate, the post supported by the base and movable with the plate between a retracted position and an extended position. The build material escapement assembly further includes a drive assembly coupled to the post to move the post between the retracted position and the extended position.
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 64/232 - Driving means for motion along the axis orthogonal to the plane of a layer
B33Y 30/00 - Apparatus for additive manufacturingDetails thereof or accessories therefor
9.
CERAMIC MATRIX COMPOSITE FASTENERS AND FASTENER SYSTEMS AND METHODS OF FORMING CERAMINC COMPOSITE MATRIX (CMC) FASTENERS AND FASTENER SYSTEMS
A ceramic matrix composite (CMC) fastener is provided and comprises a CMC material forming a body portion including a plurality of continuous fibers disposed in a matrix. The body portion defines a body length in an axial direction. The body portion defines a first end and a second end opposite the first end along the body length. The body portion includes a head formed at the first end and a shank extending from the head along the body length. The plurality of continuous fibers extend across the head and the shank.
Ceramic matrix composite (CMC) fasteners and fastener systems, as well as methods of forming CMC fasteners and fastener systems, are provided. For example, a CMC fastener includes a CMC material forming a body portion having a first plurality of continuous fibers disposed in a first matrix, and a thread portion having a second plurality of continuous fibers disposed in a second matrix. The body portion has a body length in an axial direction. The thread portion is wound in a helix about the body portion such that the second plurality of continuous fibers wrap about the first plurality of continuous fibers along the axial direction. A CMC fastener system can include the CMC fastener and a CMC threaded nut.
An auto-visual data processing and insight generation system for an aircraft engine includes a processor and a memory including processor executable instructions that cause the system to: perform video processing on image frames, wherein the video processing includes selecting frames of the image frames as key frames based upon feature criteria including a change in pixels, a trend change in graphs, or a specified filter in the frames; perform an extraction of a feature included in the key frames, wherein the feature is extracted from a two-dimensional image to generate a three-dimensional model; identify an insight regarding the feature based on the three-dimensional model to predict performance of a component of the aircraft engine; verify at least one of an analytical or physical growth model based on the identified insight; and tune a design of the aircraft engine based on a verification of the analytical or physical growth model.
A directed energy deposition (DED) additive manufacturing system for manufacturing a component from a material includes a deposition assembly having a deposition head through which material is deposited to form a top surface of a component. The top surface defines a width (w) measured in millimeters (mm). The deposition assembly also includes compression rig having a compression head. The compression head is configured to apply a compressive force (F) measured in kilonewtons (kN) to the top surface of the component. The compression rig also defines a target load (Y).
A turbofan engine (100) for an aircraft (10) includes a core cowl (118), a nacelle assembly (150) positioned radially outward of the core cowl (118) defining a bypass airflow passage (156) between the core cowl (118) and the nacelle assembly (150) where the bypass airflow passage (156) has a fan exit nozzle (158). The nacelle assembly (150) includes a fan cowl (204), a transcowl (206) positioned aft of the fan cowl (204), and a thrust reverser assembly (200). An actuation assembly (212) is operably connected to at least one of the transcowl (206) or the thrust reverser assembly (200) and is actuatable to move the transcowl (206) aft from a first position where the cascade assembly (208) is covered to a second position where the cascade assembly (208) is uncovered. The actuation assembly (212) is further actuatable to move the transcowl (206) forward from the first position to a third position to reduce an area of the fan exit nozzle (158).
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 1/09 - Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
A gas turbine engine including a compressor section, a combustion section, and a turbine section in serial flow arrangement, with the combustion section. The combustion section includes a combustion chamber and a fuel-air mixer fluidly coupled with the combustion chamber. The fuel-air mixer includes an outer wall, a center body disposed radially inward of the outer wall, a first splitter including a first lobed trailing edge, a second splitter including a second lobed trailing edge axially offset from the first lobed trailing edge, a first swirler disposed at least partially between the first splitter and the second splitter; and a set of fuel orifices located at the first lobed trailing edge, the second lobed trailing edge, or the first lobed trailing edge and the second lobed trailing edge.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F23R 3/12 - Air inlet arrangements for primary air inducing a vortex
F23R 3/14 - Air inlet arrangements for primary air inducing a vortex by using swirl vanes
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
15.
Gas Turbine Engine with Forward Swept Outlet Guide Vanes
A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan; a turbomachine drivingly coupled to the fan, the turbomachine comprising an outer casing; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip; and an accessory gearbox positioned at least partially inward of the outer casing of the turbomachine.
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 7/32 - Arrangement, mounting, or driving, of auxiliaries
F02K 3/075 - 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 controlling flow ratio between flows
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. The fuel nozzle can include both a primary and secondary fuel passage, and an additional air passage to provide for greater flame control, fuel provision, or local fuel and air mixing prior to combustion.
A turbine engine having a compressor section, combustion section, and turbine section in serial flow arrangement. The combustion section has a fuel source, an air source, and a fuel mixer assembly. The fuel mixer assembly includes a mixing tube body at least partially defining a mixing channel. The mixing tube body includes a first set of fuel passages and a second set of fuel passages fluidly coupling the fuel source to the mixing channel, wherein the first set of fuel passages supplies fuel to the mixing channel at low power conditions and the second set of fuel passages supplies fuel to the mixing channel at high power conditions.
A combustion section for a turbine engine, the combustion section comprising a wall at least partially forming a combustion chamber and a fuel nozzle extending through a respective portion of the wall, the fuel nozzle having a fuel nozzle body defining a central channel and a fuel nozzle centerline, the central channel opening to the combustion chamber at a fuel nozzle outlet, a swirler provided along the fuel nozzle body and extending into the central channel, the swirler having a helical vane wrapped circumferentially about the fuel nozzle centerline.
A combustion section for a turbine engine. The combustion section has a wall and a fuel nozzle. The wall at least partially forms a combustion chamber. The fuel nozzle opens to the combustion chamber through the wall. The fuel nozzle has a fuel nozzle body and a swirler. The fuel nozzle body defines a central channel. The swirler is provided within the central channel.
A combustion section for a turbine engine. The combustion section has a wall and a fuel nozzle. The wall at least partially forms a combustion chamber. The fuel nozzle opens to the combustion chamber through the wall. The fuel nozzle has a fuel nozzle body and a swirler. The fuel nozzle body defines a central channel. The swirler is provided within the central channel.
Apparatuses and systems are provided herein for unducted propulsion systems. The system includes an aft housing for low drag for high subsonic sustained flight. A plurality of blades are affixed to the aft housing, wherein the housing defines a flowpath curve extending from the axial extent of the aft blade root to the aft end of the aft housing. The flowpath curve is described by an axial direction parallel to an axis of rotation and a radius from the axis of rotation. The flowpath curve includes first point having a first radius where the radius reaches a maximum aft of the aft blade root and a second point forward of the first point having a second radius where the radius stops decreasing. The ratio of the first radius to the second radius is greater than or equal to 1.081.
F02C 3/067 - 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 having counter-rotating rotors
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
22.
GAS TURBINE ENGINE WITH FORWARD SWEPT OUTLET GUIDE VANES
A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.
F02K 3/068 - 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 being characterised by a short axial length relative to diameter
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/04 - Air intakes for gas-turbine plants or jet-propulsion plants
F02C 7/045 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
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
F02K 3/04 - 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
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 29/66 - Combating cavitation, whirls, noise, vibration, or the likeBalancing
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
An air-oil separation system for an oil sump in a gas turbine engine includes (a) an air-oil separator arranged to separate an air-oil mixture into an oil component and into an air component, (b) at least one air-oil inlet portion arranged to input the air-oil mixture from an oil sump into an separation chamber of the air-oil separator, the air-oil inlet portion including an inlet heat exchanger portion having an inlet coolant flow passage arranged to provide a flow of a coolant therethrough to cool the air-oil mixture input into the air-oil inlet portion, (c) at least one oil outlet arranged to provide the oil component to flow from the separation chamber of the air-oil separator to the oil sump, and (d) an air outlet arranged to provide the air component to flow out of the separation chamber of the air-oil separator.
A three-dimensional woven fabric for a composite airfoil for a turbine engine and methods of manufacturing such a fabric. The method includes forming, during weaving a plurality of reinforcing fiber tows, a spar section of the woven fabric. The method also includes forming, during weaving the plurality of reinforcing fiber tows, a first edge section integrally woven with the spar section and extending in a second direction therefrom and forming, during weaving the plurality of reinforcing fiber tows, a second edge section integrally woven with the spar section and extending in the second direction therefrom, the second edge section being positioned opposite the first edge section with a gap formed therebetween.
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/02 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising combinations of reinforcements and fillers incorporated in matrix material, forming one or more layers, with or without non-reinforced or non-filled layers
B29C 70/30 - 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
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
C04B 35/626 - Preparing or treating the powders individually or as batches
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
C08J 5/24 - Impregnating materials with prepolymers which can be polymerised in situ, e.g. manufacture of prepregs
C22C 47/06 - Pretreatment of the fibres or filaments by forming the fibres or filaments into a preformed structure, e.g. using a temporary binder to form a mat-like element
D03D 25/00 - Woven fabrics not otherwise provided for
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
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.
An airfoil assembly for turbine engines is provided. The airfoil assembly includes a plurality of airfoils and an airfoil includes a first surface defining a pressure side and a second surface defining a suction side, a plurality of first chord sections defining at least one first chord length, a plurality of second chord sections defining at least one second chord length, and a plurality of wave-shaped projections extending along the trailing edge. The plurality of wave-shaped projections each define a wave peak and a wave trough. The wave-shaped projections have an amplitude that is a proportion of an average local chord length of the airfoil. The wave-shaped projections may be designed according to a relationship to power coefficient, fan pressure ratio, and/or thrust coefficient of a turbine engine to produce an airfoil with a sculpted trailing edge feature that reduces noise during operation of the turbine engine.
An assembly for measuring impedance of a battery includes a first current loop including a first capacitor and a second capacitor in series with the first capacitor, a second current loop in parallel with the first current loop, the second current loop including a third capacitor, an impedance sensor in electrical communication with the first and second current loops, and a controller module configured to determine an impedance of the battery based on data from the impedance sensor indicating a first current frequency from the first current loop and a second current frequency from the second current loop.
G01R 31/392 - Determining battery ageing or deterioration, e.g. state of health
B60R 16/033 - Electric or fluid circuits specially adapted for vehicles and not otherwise provided forArrangement of elements of electric or fluid circuits specially adapted for vehicles and not otherwise provided for electric for supply of electrical power to vehicle subsystems characterised by the use of electrical cells or batteries
G01R 1/20 - Modifications of basic electric elements for use in electric measuring instrumentsStructural combinations of such elements with such instruments
G01R 31/389 - Measuring internal impedance, internal conductance or related variables
28.
ASPIRATING FACE SEAL ASSEMBLY FOR A ROTARY MACHINE
General Electric Deutschland Holding GmbH (Germany)
Inventor
Jalan, Prateek
Johnson, Steven Douglas
Bidkar, Rahul Anil
Portune, Grant Robert
Singh, Tajinder
Valencia, Antonio Guijarro
Abstract
An aspirating face seal assembly includes a stationary component and a seal body defining a seal face. The seal body defines a fluid feed passage within the seal body. The fluid feed passage includes an inlet, and one or more feed ports defined along the seal face. The seal body is moveably coupled to the stationary component via a resilient member. A first-pressure plenum is defined radially outward from the seal body and a second-pressure plenum is defined radially inward from the seal body. The seal assembly further includes a plenum seal forming a seal between a portion of the seal body and the stationary component and at least partially defining a third-pressure plenum in fluid communication with a high-pressure fluid source and with the inlet of the fluid feed passage. The third-pressure plenum is pressurized at a higher pressure than both the first-pressure plenum and the second-pressure plenum.
Methods of manufacturing ceramic articles are presented. For example, a method of manufacturing a ceramic article may include layering a ceramic material in a plurality of layers along a build direction, the plurality of layers forming a green component; firing the green component at a first temperature to form a porous sintered component; infiltrating the porous sintered component with a preceramic material to form an infiltrated component; and firing the infiltrated component at a second temperature to form the ceramic article.
C04B 35/18 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on silicates other than clay rich in aluminium oxide
A control circuit accesses first time-dispersed output from a first circuit and identifies a potential fault for that first circuit as a function, at least in part, of circuit output intermittency. These teachings will accommodate a variety of such circuits including, for example, a sensor (such as an aircraft-mounted jet turbine engine sensor) as well as non-sensor circuits (such as a full authority digital engine control).
An additive manufacturing apparatus includes a support plate defining a window and a resin support configured to support an uncured layer of resin. A stage is configured to hold one or more cured layers of the resin to form a component positioned opposite a support plate. A radiant energy device is positioned on an opposite side of the resin support from the stage and is operable to generate and project radiant energy in a patterned image through the window. The stage is configured to move simultaneously with the resin support from a first position to a second position in an X-axis direction.
B29C 64/124 - Processes of additive manufacturing using only liquids or viscous materials, e.g. depositing a continuous bead of viscous material using layers of liquid which are selectively solidified
B29C 64/232 - Driving means for motion along the axis orthogonal to the plane of a layer
B29C 64/236 - Driving means for motion in a direction within the plane of a layer
A method for inspecting an object includes determining a first inspection package that includes a first inspection image of the object and a first designation. The method includes determining data indicative of a second inspection package that includes a second inspection image of the object and a second designation. The method includes determining a first property of the object based on the first inspection image of the object, one or more properties maps of the object, and the first designation. The method includes determining a second property of the object based on the second inspection image of the object, the one or more properties maps of the object, and the second designation. The method includes displaying the first property and the second property or displaying data indicative of a comparison of the first property with the second property.
G06T 7/564 - Depth or shape recovery from multiple images from contours
G01B 11/25 - Measuring arrangements characterised by the use of optical techniques for measuring contours or curvatures by projecting a pattern, e.g. moiré fringes, on the object
G06T 17/20 - Wire-frame description, e.g. polygonalisation or tessellation
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
F
B
×
D
F
T
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
F
B
×
D
F
T
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 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 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 coating reinforcing fiber of a composite component is provided, The system includes a frame including at least one contact location for contacting the reinforcing fiber and a movement mechanism including an actuator. The movement mechanism is operably coupled to the frame to induce movement of the reinforcing fiber relative to the frame. Methods are also provided for coating such a fiber.
B28B 11/04 - Apparatus or processes for treating or working the shaped articles for coating
B28B 19/00 - Machines or methods for applying the material to surfaces to form a permanent layer thereon
B28B 23/04 - Arrangements specially adapted for the production of shaped articles with elements wholly or partly embedded in the moulding material wherein the elements are reinforcing members the elements being stressed
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 turbine engine comprising: a compressor section, a combustion section, and a turbine section in serial flow arrangement, with the combustion section comprising a combustion chamber and a fuel nozzle assembly fluidly coupled with the combustion chamber at a fuel nozzle assembly outlet, wherein the fuel nozzle assembly comprises a gaseous fuel passage defined by an inner surface of a centerbody, a plurality of annularly shaped air supply passages circumscribing the centerbody including at least a first air supply passage and a second air supply passage circumscribing the first air supply passage, at least one splitter having an annular shape circumscribing the centerbody and separating adjacent air supply passages of the plurality of air supply passages, a fuel nozzle body circumscribing the plurality of air supply passages and the at least one splitter and at least one orifice plate disposed within one of the plurality of the air supply passages.
A gas turbine engine, comprising: a compressor section, a combustion section, and a turbine section, with the combustion section including: a flow turning combustor comprising: a combustor liner that at least partially defines a combustion chamber including: a combustor inlet having an inlet height, a combustor outlet having an outlet height and fluidly coupled with the turbine section; a primary section extending from the combustor inlet and including a primary length; and a transition section extending from the primary section to the combustor outlet and defining a turn; a hydrogen fuel supply; and a fuel nozzle assembly fluidly coupled with the hydrogen fuel supply and the combustor inlet.
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
Rotatable trunnions for variable pitch blades are disclosed herein. An variable pitch blade assembly comprises a trunnion defining a pitch axis, the trunnion including: a first sleeve defining a face of the trunnion; and a second sleeve within the first sleeve, the second sleeve including a pedestal and a column extending from the pedestal; a first spar for a fan blade, the first spar having a base portion and a mounting portion, the mounting portion protruding from the face of the trunnion, the base portion including an inner surface defining a bore, wherein the column of the second sleeve is disposed within the bore of the base portion, wherein the first sleeve, the second sleeve, and the first spar are coaxial with the pitch axis; and a second spar extending from the face of the trunnion, the second spar laterally offset relative to the first spar.
A gas turbine engine has a turbomachine comprising compressor, combustion, and turbine sections. The gas turbine engine defines a maximum exhaust gas temperature, a maximum drive turbine shaft torque, and a corrected specific power. The gas turbine engine includes a propeller and an air inlet. The propeller has a blade that defines a blade path area when rotated about a propeller axis. The blade path area has a radius. The air inlet defines an opening wherein a radial distance is defined from the propeller axis to a location within the opening, wherein the radial distance is equal to or less than 60% of the radius of the blade path.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02C 6/06 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
An airfoil assembly has an airfoil portion. The airfoil portion extends between a root and a tip in a spanwise direction a span length. The airfoil portion extends between a leading edge and a trailing edge in a chordwise direction a chord length. The airfoil assembly includes a cap overlaying a portion of the outer wall. The cap extends between a cap fore edge and a cap aft edge a cap chord length. The cap extends between a cap root and a cap tip a cap span length.
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. 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. The composite material includes a fiber preform architecture that includes a plurality of fiber bands. Each band of the plurality of fiber bands is placed one at a time, at a predetermined position and orientation, to generate an interwoven and interlocking pattern between the fiber bands.
A gas turbine engine includes: a turbomachine including 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 gearbox mechanically coupling the drive turbine of the turbomachine to the fan. The 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. The engine includes a gearbox efficiency rating of 0.10-1.8 or an overall engine efficiency rating of 0.57-8.0.
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 plurality of variable pitch fan blades formed of a composite material, each 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. In certain examples, the engine further includes a pitch change mechanism.
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.
An inspection system and related methods are provided. The inspection system includes an inspection camera, a plurality of light sources collocated with the inspection camera, and a post processing system. The plurality of light sources output directional light that have different respective ranges of light wavelengths. The inspection camera is configured to capture image data while a surface of interest is being illuminated with the directional light. Further, the post processing system is configured to receive the image data, process portions of the image data into a plurality of images that include distinct images corresponding to the different respective ranges of light wavelengths. The plurality of images can be reviewed to identify an abnormal region of the surface of interest.
A method of heat-treating an additively-manufactured ferromagnetic component is presented and a related ferromagnetic component is presented. A saturation flux density of a heat-treated ferromagnetic component is greater than a saturation flux density of an as-formed ferromagnetic component. The heat-treated ferromagnetic component is further characterized by a plurality of grains such that at least 25% of the plurality of grains have a median grain size less than 10 microns and 25% of the plurality of grains have a median grain size greater than 25 microns.
C21D 9/02 - Heat treatment, e.g. annealing, hardening, quenching or tempering, adapted for particular articlesFurnaces therefor for springs
C22C 19/07 - Alloys based on nickel or cobalt based on cobalt
C22C 33/02 - Making ferrous alloys by powder metallurgy
C22C 38/10 - Ferrous alloys, e.g. steel alloys containing cobalt
H01F 1/20 - Magnets or magnetic bodies characterised by the magnetic materials thereforSelection of materials for their magnetic properties of inorganic materials characterised by their coercivity of soft-magnetic materials metals or alloys in the form of particles, e.g. powder
H01F 41/04 - Apparatus or processes specially adapted for manufacturing or assembling magnets, inductances or transformersApparatus or processes specially adapted for manufacturing materials characterised by their magnetic properties for manufacturing cores, coils or magnets for manufacturing coils
H02K 7/00 - Arrangements for handling mechanical energy structurally associated with dynamo-electric machines, e.g. structural association with mechanical driving motors or auxiliary dynamo-electric machines
H02K 15/00 - Processes or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines
H02K 15/02 - Processes or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of stator or rotor bodies
H02K 15/12 - Impregnating, moulding insulation, heating or drying of windings, stators, rotors or machines
Secondat De Montesquieu, Antoine Claude Baudoin Raoul Marie
Soulat, Laurent
Schvallinger, Mickaël Franck Antoine
Abstract
A triple-flow aircraft turbine engine, having two coaxial annular walls, rotor blading an annular separator arranged downstream of the rotor blading and between the two walls, and having, upstream, an annular nose, stationary guide vanes connected to the nose, and variable-pitch guide vanes downstream of the stationary guide vanes.
F02K 3/077 - 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 the plant being of the multiple flow type, i.e. having three or more flows
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
An airfoil assembly has an airfoil portion. The airfoil portion extends between a root and a tip in a spanwise direction a span length. The airfoil portion extends between a leading edge and a trailing edge in a chordwise direction a chord length. The airfoil assembly includes a cap overlaying a portion of the outer wall. The cap extends between a cap fore edge and a cap aft edge a cap chord length. The cap extends between a cap root and a cap tip a cap span length.
An airfoil assembly has an airfoil portion with an outer wall bounding an interior. The outer wall extends between a root and a tip in a spanwise direction. The outer wall extends between a leading edge and a trailing edge in a chordwise direction. The outer wall includes a cap overlaying a portion of the outer wall. The cap has an exterior surface and corrugations provided along the exterior surface.
A hybrid engine system includes a duct, a turbine engine disposed within the duct, the turbine engine including a fan section, a compressor section, a combustion section, a turbine section, and a shaft coupled to the compressor section, an electrical power generator mechanically coupled to the shaft, the electrical power generator being configured to convert at least a portion of mechanical power generated by the turbine engine into electrical power, and a part-span inlet guide vane disposed upstream of the turbine engine within the duct, the part-span inlet guide vane including inlet guide vanes that are rotatable to control an amount of airflow and a direction of the airflow towards the fan section of the turbine engine.
A blended wing aircraft is provided, defining a longitudinal direction and a lateral direction, the blended wing aircraft including a body; a pair of wings extending outward from the body along the lateral direction; and a propulsion system comprising an engine mounted to the body, the engine defining an axial direction and having a combustion section and a fan, the fan positioned downstream of the combustion section along the axial direction.
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 variable pitch primary fan driven by the turbomachine; a secondary fan located downstream of the primary fan within the inlet duct; and a fluid transfer system for supplying fluid to the primary fan. 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
55.
COATED COMPONENTS FOR COKE ABATEMENT IN GAS TURBINE ENGINES
A coated component for coke abatement in a gas turbine engine. The coated component includes a metal substrate defining, at least in part, a flow passage for a hydrocarbon fluid, and a nanophase separated catalytic coating deposited on the metal substrate to be exposed to the flow passage for abating coke formation from the hydrocarbon fluid. The nanophase separated catalytic coating includes a substantially pure transition metal phase, a substantially pure noble metal phase, and a substantially pure transition metal oxide phase.
A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
(
T
OUT
A
DTExit
)
2
*
EGT
A
HPCExit
*
10
-
11
;
A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
(
T
OUT
A
DTExit
)
2
*
EGT
A
HPCExit
*
10
-
11
;
wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02C 6/06 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
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 unducted fan propulsor may further include an outlet nozzle that expels an exhaust stream at a non-zero angle with the centerline of the unducted fan propulsor such that the centerline is oriented downwardly relative to the exhaust stream. The outlet nozzle may further include a core cowl shaped to cause a bypass or third stream flow to entrain a core exhaust stream.
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 gas turbine engine is provided. The gas turbine engine includes a turbomachine having a turbomachine having a compressor section, a combustor, and a turbine section arranged in serial flow order, the turbomachine further including an outer casing; and a fuel cell assembly positioned within the outer casing of the turbomachine, the fuel cell assembly including a fuel cell, an inlet line in fluid communication with an inlet of the fuel cell, and an output products line in fluid communication with an outlet of the fuel cell for receiving output products from the fuel cell, wherein the inlet line is positioned to be in thermal communication with the output products during operation of the gas turbine engine.
F02C 7/232 - Fuel valvesDraining valves or systems
H01M 8/04014 - Heat exchange using gaseous fluidsHeat exchange by combustion of reactants
H01M 8/04089 - Arrangements for control of reactant parameters, e.g. pressure or concentration of gaseous reactants
H01M 8/04111 - Arrangements for control of reactant parameters, e.g. pressure or concentration of gaseous reactants using a compressor turbine assembly
H01M 8/0612 - Combination of fuel cells with means for production of reactants or for treatment of residues with means for production of gaseous reactants from carbon-containing material
H01M 8/12 - Fuel cells with solid electrolytes operating at high temperature, e.g. with stabilised ZrO2 electrolyte
A combustion liner assembly includes a metal liner, a ceramic matrix composite (CMC) liner tile, and a fastener. The metal liner defines a liner opening extending through a hot side and a cold side of the metal liner. The CMC liner tile includes first and second liner tile connection members each extending from the CMC liner tile and through the liner opening from the hot side to the cold side of the metal liner. The fastener is disposed on the cold side of the metal liner and couples the first and second liner tile connection members, thus coupling the CMC liner tile to the metal liner. The fastener being configured, during operation of the turbine engine, to allow radial movement and circumferential movement of the CMC liner tile relative to the metal liner to accommodate thermal expansion of the CMC liner tile.
Secondat De Montesquieu, Antoine Claude Baudouin Raoul Marie
Soulat, Laurent
Abstract
An aircraft propulsion assembly having a nacelle surrounding a three-flow turbomachine including a gas generator, a fan that accelerates an air flow through the nacelle, an annular element between the generator and the nacelle defining a first duct and a second duct. The annular element can have a nose for splitting the flow into an air flow through the first duct and into an air flow through the second duct, the assembly can include a stator vane mounted between the nose and the fan, and a stator vane between the generator and the annular element, mounted between the nose and a rotor vane of a compressor of the generator, or between the annular element and the nacelle.
F02K 3/077 - 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 the plant being of the multiple flow type, i.e. having three or more flows
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 29/00 - Power-plant nacelles, fairings or cowlings
B64D 35/00 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
62.
Fuel vaporization power turbine engine, method of assembly and method of use
Some embodiments provide engines comprising: a fuel line; a gas combustion system; a first power turbine positioned within an exhaust gas stream and configured to operate at a first rotation speed; a second power turbine downstream from the first power turbine, and configured to operate at a second rotation speed that is less than the first rotation speed; and a heat exchanger configured to receive the exhaust gas stream, wherein a portion of a fuel line is positioned to receive heat from the heat exchanger configured to heat the fuel; a gear ratio (GR) is defined by a ratio of the second rotation speed to the first rotation speed; a combination of the first power turbine and the second power turbine includes a number of blade rows (NBR); and a system effectiveness control (SEC) is established as a function of the GR and the NBR.
F01D 1/26 - Non-positive-displacement machines or engines, e.g. steam turbines characterised by counter-rotating rotors subjected to same working-fluid stream without intermediate stator blades or the like traversed by the working-fluid substantially axially
A gas turbine engine includes a compressor section, a combustion section defining a combustion chamber, and a turbine section disposed in serial flow order along a central axis of the gas turbine engine. The combustion section includes a plurality of fuel nozzles in fluid communication with the combustion chamber. The plurality of fuel nozzles define a fuel nozzle pitch extending between a fuel nozzle centerline of adjacent ones of the plurality of fuel nozzles. The turbine section includes a plurality of vanes and a plurality of rotor blades. The turbine section defines a clocking pitch fraction in degrees about the central axis and the clocking pitch fraction is defined between a peak temperature region and a midpoint of a vane pitch. The vane pitch is defined between the leading edge of adjacent ones of the plurality of vanes.
A squeeze film damper assembly for a turbine engine includes an annular bearing support and an annular damper housing. The annular bearing support includes an inner support segment and an outer support segment located radially outward of the inner support segment. The inner support segment and the outer support segment are spaced apart to define a support channel therebetween. The annular damper housing is at least partially received in the support channel to define an inner damping chamber and an outer damping chamber, the inner damping chamber being radially inward of the outer damping chamber, the inner damping chamber and the outer damping chamber each capable of being filled with an amount of lubricant to provide a squeeze film damper at the inner damping chamber, the outer damping chamber, or both.
In one aspect, a blended wing aircraft includes a blended wing body and an engine including a nacelle defining an inlet and an outlet. The engine further includes a fan positioned within the nacelle between the inlet and the outlet, with the engine being supported relative to the blended wing body such that the inlet of the nacelle is configured to receive boundary layer air from the blended wing body. In addition, the blended wing aircraft includes a plurality of surface features positioned upstream of the fan, with each surface feature of the plurality of surface features comprising a projection extending outwardly from an adjacent surface of the blended wing aircraft such that the projection is configured to interact with the boundary layer air.
An apparatus to reduce vibration of seal segments for a gas turbine engine is disclosed. An example apparatus includes a rotor, a stator, and a radial seal positioned between the rotor and the stator, the radial seal including a plurality of seal segments spanning a circumference of the rotor, a quantity of the plurality of seal segments corresponding to a prime number.
A turbine engine has a compressor section, a combustion section, and a turbine section in serial flow arrangement. The combustion section has a combustor liner and dome wall collectively forming at least a portion of a combustion chamber. The dome wall has a fuel nozzle opening. The combustion section has a fuel nozzle assembly extending through the fuel nozzle opening. The fuel nozzle assembly has a first body, a second body, a first swirler and a second swirler.
A snake-arm robot and a servicing device are mechanically coupled. The mechanical coupling is accomplished by a longitudinal insertion of the snake-arm robot into the servicing device or the servicing device into the snake-arm robot. An actuator moves the snake-arm robot through a passage within an engine until the snake-arm robot reaches a desired location. The movement of the snake-arm robot concurrently moves the servicing device through the passage. Subsequently, the snake-arm robot is de-coupled from the servicing device and the snake-arm robot is removed from the engine while leaving the servicing device in place within the engine.
A method for repairing a component that comprises a ceramic matrix composite (“CMC”) material includes forming a repair insert defined by a repair geometry where the repair geometry is based on a repair area of the component, and the repair insert comprises a monolithic ceramic. Inserting the repair insert into the repair area and applying a CMC face sheet to the repair insert. The method further includes bonding the repair insert to the CMC face sheet, the repair insert to the component, and the CMC face sheet to the component. The method also includes thermally processing and densifying at least one of the repair insert or the CMC face sheet in the repair area.
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 gas turbine engine includes a drive turbine, a fan, and a gearbox. The fan has a 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 gearbox mechanically couples the drive turbine to the fan. 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. The drive turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage divided by the annular exit area of a forward-most rotating stage. In some instances, the area ratio is within a range of 2.0-6.5. Additionally (or alternatively), the low-pressure turbine includes an area-EGT ratio within a range of 1.05-1.6.
A method for additively manufacturing three-dimensional objects includes generating a laser beam with a laser beam source and splitting the laser beam to form a plurality of beamlets. The plurality of beamlets are collimated via an optical device. The method also includes independently controlling respective beamlets of the plurality of beamlets via respective channels of a multi-channel optical modulator disposed downstream of the optical device to at least one of steer or modulate the respective beamlets. A scanning device disposed downstream of the multi-channel optical modulator scans the respective beamlets over at least a portion of a target plane.
B23K 26/067 - Dividing the beam into multiple beams, e.g. multi-focusing
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]
B29C 64/282 - Arrangements for irradiation using multiple radiation means, e.g. micromirrors or multiple light-emitting diodes [LED] of the same type, e.g. using different energy levels
A CVFDR system of an aircraft includes a cockpit voice and flight data recorder (CVFDR) communicatively coupled, via a data communication network, to a set of flight recorder modules. The CVFDR receives a first voltage from a remote first power source. In the event of an interruption of the first voltage, the CVFDR receives a second voltage from a local second power source for a predetermined period.
B64D 45/00 - Aircraft indicators or protectors not otherwise provided for
H04L 67/12 - Protocols specially adapted for proprietary or special-purpose networking environments, e.g. medical networks, sensor networks, networks in vehicles or remote metering networks
A method for inspecting an object includes receiving or determining inspection image data, the inspection image data including an inspection image pixel array with at least one inspection image pixel in the inspection image pixel array having a pixel property associated therewith. The method includes receiving via a processor a user input associated with a continuous segment of inspection image pixels in the inspection image pixel array. The method includes determining a property of the object based on the pixel properties associated with the continuous segment of inspection image pixels in the inspection image pixel array.
A gas turbine engine includes a turbomachine comprising a turbomachine 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 speed reduction device mechanically coupling the turbomachine to the fan; wherein the gas turbine engine defines a fan leading edge to trailing edge compression factor (FLTCF) or a fan leading edge to trailing edge opening ratio (FLTOR). The FLTCF is greater than or equal to 1.05 and less than or equal to 1.8.
A gearbox assembly includes a gearbox having a gear assembly 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 is defined by
A gearbox assembly includes a gearbox having a gear assembly 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 is defined by
V
G
V
GB
.
A gearbox assembly includes a gearbox having a gear assembly 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 is defined by
V
G
V
GB
.
VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly and a lubrication system. The lubrication system includes a sump that is a primary reservoir having a first lubricant level and a secondary reservoir in the gearbox assembly. The secondary reservoir has a second lubricant level. The lubrication system fills the secondary reservoir with a lubricant between the first lubricant level and the second lubricant level. The gear assembly collects the lubricant in the secondary reservoir to supply the lubricant to the gear assembly.
The present invention relates to improved radiopharmaceutical compositions in sealed containers, where the container closure has an ETFE (ethylene-tetrafluoroethylene copolymer) coating. Also disclosed are kits for radiopharmaceutical preparation using the sealed containers, as well as methods of preparation of radiopharmaceuticals using the sealed containers.
A61K 51/12 - Preparations containing radioactive substances for use in therapy or testing in vivo characterised by a special physical form, e.g. emulsion, microcapsules, liposomes
A gas turbine engine, including an accessory system cooling system comprising a cooling system inlet and a duct in fluid communication with the cooling system inlet; and a turbomachine comprising a compressor section, a combustion section, and a turbine section, the turbomachine defining a working gas flowpath and further comprising a cooled cooling air (CCA) system, the CCA system comprising a cold side bleed assembly and a CCA heat exchanger in thermal communication with the cold side bleed assembly, wherein the cold side bleed assembly defines an inlet in fluid communication with the duct of the accessory system cooling system at a location downstream of the cooling system inlet.
A turbine engine including a turbo-engine, a gearbox assembly, a propulsor, and a lubrication system. The lubrication system includes a lubricant tank that stores lubricant therein, one or more primary gearbox lubricant supply lines in fluid communication with the lubricant tank and the gearbox assembly, one or more secondary gearbox lubricant supply lines in fluid communication with the lubricant tank and the gearbox assembly, and a lubricant pump for supplying the lubricant to the gearbox assembly from the lubricant tank through the one or more primary gearbox lubricant supply lines and the one or more secondary gearbox lubricant supply lines. The lubrication system modulates a mass flow rate of the lubricant to the gearbox assembly through at least one of the one or more primary gearbox lubricant supply lines or the one or more secondary gearbox lubricant supply lines.
Gas turbine engines including thermal management systems with recirculating fuel circuits and related methods are disclosed herein. An example gas turbine engine disclosed herein includes a fuel flow line, a fuel inlet, a combustor fluidly coupled to the fuel inlet by the fuel flow line, and a recirculation flow circuit coupled to the fuel flow line between the fuel inlet and the combustor, the recirculation flow circuit including a cooler and a heat-producing component coupled to the fuel flow line downstream of the fuel inlet along the fuel flow line.
Apparatus, systems, and articles of manufacture are disclosed to dynamically support axial thrust in pumps. An example thrust bearing on a shaft disclosed herein includes a first thrust pad, a second thrust pad, and a thrust disc to be rigidly coupled to the shaft at a first end, the thrust disc positioned between the first and second thrust pads, the thrust disc including a tapered edge at a second end opposite the first end.
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F01D 3/04 - Machines or engines with axial-thrust balancing effected by working fluid axial thrust being compensated by thrust-balancing dummy piston or the like
F01D 15/08 - Adaptations for driving, or combinations with, pumps
A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT/(AHPCExit2×1000).
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
A combustion section for a turbine engine. The combustion section has a combustor and a fuel nozzle. The combustor has a combustion chamber. The fuel nozzle has a body, and a vane. The body defines a centerline. The body has a central channel. The central channel has a compressed air flow passage and a mixer. The vane extends from the body. The vane has an outer wall. The vane has a plurality of lobes.
Methods, systems, and apparatus are disclosed to provide a pressurized fluid to components of a fluid pump. An example flow control system to provide a pressurized lubricant to a secondary flow network disposed within a fluid pump includes sensors to measure parameters of a fluid corresponding to fluid flow; a recirculation loop fluidly coupled to a secondary inlet of the pump, the recirculation loop to provide a first flowpath, wherein the secondary inlet is an inlet to the secondary flow network; a bypass circuit fluidly coupled to the secondary inlet to provide a second flowpath; and a controller to direct the fluid flow to the first flowpath or the second flowpath based on sensor data from the sensor, the sensor data indicative of a state of the fluid.
Systems, apparatus, articles of manufacture, and methods to provide for cross-diffuser bleed are provided herein. An example gas turbine engine includes a frame defining a cavity in a forward side of a diffuser of the gas turbine engine; a compressor including the diffuser, the diffuser defining a primary flow path to provide air flow to a combustor and including at least one conduit, the at least one conduit fluidly coupled to the cavity; and a downstream sink fluidly coupled to the cavity via the at least one conduit defining at least a portion of a bleed air path.
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
Example apparatus, systems, and methods for rapid active clearance control of inter-stage and mid-stage seals are disclosed. An example apparatus to control clearance for a turbine engine comprises a case surrounding at least part of the turbine engine and defining an opening therethrough; a nozzle, the nozzle including a reference pressure sensor and a static pressure sensor on a tip of the nozzle; an actuator including a multilayer stack of material, a rod coupled to the first actuator and coupled to the nozzle through the opening in the case, the rod to move the nozzle based on contraction or expansion of the multilayer stack of material; and a controller to calculate and set the clearance between the rotor and the nozzle by supplying an electrical current to the multilayer stack to cause the multilayer stack to at least one of expand or contract.
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 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 recoater assembly is movable across a working surface in a forward stroke and a return stroke and includes a powder plow. One or more actuators are coupled to the powder plow and actuable to provide vertical movement of the powder plow with respect to the working surface. A controller is operable to control the one or more actuators to maintain the powder plow in a retracted position during the forward stroke, lower the powder plow to an extended position when the powder plow is at or near a return area, enable the powder plow to be lifted from the extended position to a partially retracted position a predefined vertical distance from the working surface, and maintain the powder plow in the partially retracted position while traversing a portion of the working surface corresponding to a build platform during the return stroke.
General Electric Company Polska Sp. z o.o. (Poland)
Inventor
Sibbach, Arthur William
Pazinski, Adam Tomasz
Abstract
An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.
F02C 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
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
F02C 7/055 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with intake grids, screens or guards
Gas turbine engines and methods of their operation are provided. For example, a method of operating a gas turbine engine comprises selectively engaging and disengaging an engine clutch disposed between a low speed spool and a rotor assembly of the engine. Engagement or disengagement of the engine clutch is selected based on an operating condition of an aircraft comprising the engine. Further, an inter-spool clutch disposed between the low speed spool and a high speed spool of the engine transitions between engaged and disengaged, disengaging when the high speed spool reaches a speed greater than an operational speed of the low speed spool. Similarly, a gas turbine engine comprises an engine clutch configured to selectively position a low speed spool in operative communication with a rotor assembly and an inter-spool clutch configured to position the low speed spool in operative communication with a high speed spool.
An aircraft monitoring system includes an avionics communication bus structure, at least one network member user device that transmits a broadcast message onto the avionics communication bus structure, and at least one non-member user device that receives the broadcast message transmitted onto the avionics communication bus, processes the received broadcast message, and transmits output data to a monitoring device. The at least one non-member user device includes a bus interface, and a field programmable gate array (FPGA) that communicates with the bus interface. The FPGA is programmed to function as a main finite state machine that processes the broadcast message from the bus interface, and a transfer finite state machine that generates output data and transfers the generated output data to an output processor that communicates with the monitoring device. The monitoring device outputs a monitored data report.
An electrical component and shield assembly can include a set of conductors, with the shield assembly having a body at least partially encircling the set of conductors. The shield assembly can include a first shield layer radially spaced from a second shield layer. The first shield layer can include a magnetic material. The second shield layer can an electrically-conductive material.
A method of reshaping and reconstruction includes removing a deformed portion of a component to define a native component portion and adding a replacement portion to the native component portion. The replacement portion is adaptively machined based on one or more parameters of the native component portion and based on one or more original design parameters of the component.
F01D 5/00 - BladesBlade-carrying membersHeating, heat-insulating, cooling, or antivibration means on the blades or the members
G05B 19/401 - Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by control arrangements for measuring, e.g. calibration and initialisation, measuring workpiece for machining purposes
An airfoil assembly for a gas turbine engine includes at least two airfoils. Each airfoil of the at least two airfoils includes a leading edge, a trailing edge, a pressure side extending between the leading edge and the trailing edge, and a suction side opposite the pressure side. Each airfoil defines a chord line extending from the leading edge to the trailing edge along the pressure side, a suction side tangency point, a first reference line extending from the leading edge perpendicular to the chord line, a second reference line extending from the leading edge to the suction side tangency point, a camber line angle defined between the first reference line and the second reference line, and a solidity based on an axial width of the at least two airfoils and a pitch between adjacent ones of the at least two airfoils.
A Hybrid Turbine Starter (HTS) for a gas turbine engine. The HTS has a turbine member, a turbine shaft, an electric machine, and a controller module. The electric machine having a rotor and a stator. The stator has a set of windings. The controller module is configured to supply, during a first timeframe, a first electrical pulse. The controller module is configured to supply, during a second timeframe, a second electrical pulse.
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Sibbach, Arthur W.
Lobocki, Marcin Jacek
Bulsiewicz, Tomasz Jan
Wachulec, Marcin Krzysztof
Clements, Jeffrey D.
Abstract
Gas turbine engines with inlet guide vanes are described herein. The inlet guide vanes have throat solidity (TS), variable throat solidity (VTS), and span throat solidity (STS) values within particular ranges.
General Electric Deutschland Holding GmbH (Germany)
Inventor
Prabhakaran, Satish
Osama, Mohamed
Abstract
A vehicle includes a gas turbine engine having at least two spools and an associated power system. The power system includes two independent power subsystems, including a first power subsystem for managing power transfer between spools and a second power subsystem for supplying a base power load to the vehicle. The first power subsystem has a first electric machine mechanically coupled with a first spool of the gas turbine engine and a second electric machine mechanically coupled with a second spool. The second electric machine is electrically coupled with the first electric machine such that electrical power is transmittable therebetween. The second power subsystem has a third electric machine mechanically coupled with one of the spools. The third electric machine is electrically coupled with a load positioned offboard the gas turbine engine. The first power subsystem and the second power subsystem are electrically decoupled from one another.
F02C 3/113 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission with variable power transmission between rotors
B64D 27/02 - Aircraft characterised by the type or position of power plants
B64D 27/12 - Aircraft characterised by the type or position of power plants of gas-turbine type within, or attached to, wings
B64D 29/02 - Power-plant nacelles, fairings or cowlings associated with wings
B64D 31/18 - Power plant control systemsArrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
99.
METHODS AND APPARATUS TO REMOVE LIQUID FROM A HOUSING
Methods and apparatus to remove liquid from a housing are disclosed. An example system includes a pump including a chamber, a shaft positioned at least partially in the chamber, and a bearing to support the shaft, the chamber including a chamber inlet and a chamber outlet, the chamber to hold a fluid in a first state, a first conduit to carry the fluid in a second state of the fluid, the first conduit fluidly coupled to the chamber inlet, a second conduit to carry the fluid in the first state of the fluid, the second conduit fluidly coupled to the chamber outlet, and at least one jet pump to deliver a mixture of the fluid in the first state of the fluid and the second state of the fluid to a third conduit, and a heat exchanger coupled to the first conduit upstream of the first inlet.
A cable that includes a conductor defining a hollow interior, a casing surrounding the conductor, an electrical insulator positioned between the conductor and the casing, and a fluid positioned within the hollow interior of the conductor.
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 35/021 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants
H01B 3/16 - Insulators or insulating bodies characterised by the insulating materialsSelection of materials for their insulating or dielectric properties mainly consisting of inorganic substances gases
H01M 50/249 - MountingsSecondary casings or framesRacks, modules or packsSuspension devicesShock absorbersTransport or carrying devicesHolders specially adapted for aircraft or vehicles, e.g. cars or trains
H01M 50/298 - MountingsSecondary casings or framesRacks, modules or packsSuspension devicesShock absorbersTransport or carrying devicesHolders characterised by the wiring of battery packs
H01M 50/588 - Means for preventing undesired use or discharge for preventing incorrect connections inside or outside the batteries outside the batteries, e.g. incorrect connections of terminals or busbars
H01M 50/59 - Means for preventing undesired use or discharge for preventing incorrect connections inside or outside the batteries characterised by the protection means
