Methods and apparatus to mitigate movement between compressor casing segments are disclosed. An example compressor casing comprises a first annular casing segment, a second annular casing segment surrounding the first segment, the second segment including an opening through a first surface and a second surface of the second segment, the first surface facing the first segment, a cylindrical body extending between the first and second segments, a first end of the cylindrical body attached to the first segment, a second end of the cylindrical body positioned within the opening of the second segment, the second end of the cylindrical body held in compression within the opening, and a spring surrounding the cylindrical body, the spring aligned to a longitudinal axis of the cylindrical body, the spring positioned between the first and second annular casing segments, the spring to resist movement between the first and second annular casing segments.
A heat exchanger assembly includes an absorber element defining an axial flowpath for a fluid extending along an axis. The absorber element includes an outer support wall and one or more support members extending radially inward from the outer support wall with respect to the axis. The heat exchanger assembly also includes one or more heat exchange elements floatably coupled to the one or more support members where the one or more heat exchange elements extend along the axial flowpath.
F24S 20/20 - Solar heat collectors for receiving concentrated solar energy, e.g. receivers for solar power plants
F24S 10/75 - Solar heat collectors using working fluids the working fluids being conveyed through tubular absorbing conduits with enlarged surfaces, e.g. with protrusions or corrugations
F24S 70/16 - Details of absorbing elements characterised by the absorbing material made of ceramicDetails of absorbing elements characterised by the absorbing material made of concreteDetails of absorbing elements characterised by the absorbing material made of natural stone
F24S 70/60 - Details of absorbing elements characterised by the structure or construction
A turbine engine is provided. The gas turbine engine includes: a rotor; a stator comprising a carrier; a seal assembly disposed between the rotor and the stator, the seal assembly comprising a plurality of seal segments supported at least in part by the carrier, the plurality of seal segments having a first seal segment and a second seal segment, the first and second seal segments each having a seal face forming a fluid bearing with the rotor; and a seal support assembly comprising a tangential spring extension extending between the first seal segment and the second seal segment for biasing the first seal segment away from the second seal segment in the circumferential direction.
A heat exchanger assembly includes a first stage heat exchange section defining one or more intakes for receiving a fluid. The first stage heat exchange section includes one or more preheater elements radially insertable into the first stage heat exchange section with respect to a centrally disposed axis of the heat exchanger assembly. The preheater elements define one or more preheat flowpaths extending radially inward from the intakes. The preheater elements transfer thermal energy to or from the fluid as the fluid flows from the intakes through the preheat flowpaths. A centrally disposed second stage heat exchange section defines an axial flowpath that is fluidly connected to the preheat flowpaths to receive the fluid from the preheat flowpaths. The second stage heat exchange section transfers thermal energy to or from the fluid as the fluid flows through the axial flowpath.
A heat exchanger assembly includes a first stage heat exchange section defining one or more intakes for receiving a fluid. The first stage heat exchange section includes one or more preheater elements defining one or more preheat flowpaths extending radially inward from the one or more intakes with respect to a centrally disposed axis of the heat exchanger assembly. The one or more preheater elements include one or more guide vanes configured to guide a flow of the fluid in a spiral path from the one or more intakes toward the centrally disposed axis. A centrally disposed second stage heat exchange section is fluidly connected to the one or more preheat flowpaths to receive the fluid from the one or more preheat flowpaths.
F24S 20/20 - Solar heat collectors for receiving concentrated solar energy, e.g. receivers for solar power plants
F24S 10/75 - Solar heat collectors using working fluids the working fluids being conveyed through tubular absorbing conduits with enlarged surfaces, e.g. with protrusions or corrugations
F24S 70/60 - Details of absorbing elements characterised by the structure or construction
6.
SYSTEM AND APPARATUS FOR REDUCING BOW WAVES IN GAS TURBINE ENGINES
A gas turbine engine includes a compressor section, a combustion section including an inner liner and an outer liner spaced from the inner liner, and a turbine section. The inner liner and outer liner at least partially define a combustion chamber. The turbine section includes an inner band extending between an upstream side and a downstream side opposite the upstream side and an outer band spaced from the inner band and extending between the upstream side and the downstream. The inner band and outer band at least partially define a working gas flow path. One or both of the inner band and the outer band include a step portion adjacent the upstream side and a body portion extending from the step portion to the downstream side. The step portion extends in a radial direction past the body portion.
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
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
Example variable bleed valve assemblies for a gas turbine engine are disclosed herein. An example variable bleed valve assembly includes a port extending radially outward from a main flow path of the gas turbine engine, a door positioned at an exit of the port, and an acoustic black hole (ABH) assembly coupled to the door. The ABH assembly includes a body and a plurality of plates coupled to an interior surface of the body. The body defines a cavity having a depth. Each of the plurality of plates has a surface area, and the plurality of plates are arranged such that the surface areas of the plurality of plates vary along the depth in a radially outward direction of the gas turbine engine.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
A turbine engine includes a fan and a turbomachine defining an engine centerline. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbine engine also includes a set of composite airfoils circumferentially arranged about the engine centerline. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
A method of forming an assembly. The assembly having a stack of plies and an airfoil portion. The airfoil portion has an outer wall extending between a root and a tip in a spanwise direction, and between a leading edge and a trailing edge. The assembly has a first set of plies and a second set of plies. The first set of plies form at least a portion of the airfoil portion.
A gas turbine engine is provided having a turbomachine; a primary fan driven by the turbomachine; a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10; and an airflow mixer assembly comprising: a low pressure duct positioned within the core cowl; a plenum extending along the circumferential direction located downstream of the low pressure duct and in fluid communication with the low pressure duct; an inner shroud located downstream of the plenum and downstream of the exhaust and in fluid communication with both the plenum and the exhaust; and an outer shroud located outward of the plenum along the radial direction and extending along the circumferential direction and along the axial direction over and downstream of the inner shroud.
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
F02C 7/141 - Cooling of plants of fluids in the plant of working fluid
F02C 7/16 - Cooling of plants characterised by cooling medium
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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
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 3/06 - 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
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
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
11.
REDUCED COMMON MODE VOLTAGE PULSE WIDTH MODULATION SWITCHING SCHEME WITH CAPACITOR VOLTAGE BALANCING FOR A MULTILEVEL POWER CONVERTER
A multilevel power converter includes a plurality of switches, a first DC link capacitor, a second DC link capacitor, and one or more processors configured to: generate, for a duty cycle of the multilevel power converter, a pulse width modulated pulse pattern in accordance with a reduced common mode voltage scheme; modify the pulse width modulated pulse pattern to render a modified pulse pattern; and cause the plurality of switches to implement the duty cycle based at least in part on the modified pulse pattern to render a common mode voltage pulse to balance voltages at the first DC link capacitor and the second DC link capacitor.
H02M 7/797 - Conversion of AC power input into DC power outputConversion of DC power input into AC power output with possibility of reversal by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only
12.
AUTOMATED DE-POWDERING OF ADDITIVE MANUFACTURING BUILD
An automated de-powdering system comprises a separation station having a support structure configured to support a build box containing an additive manufacturing build. An actuator is configured to move at least a portion of the additive manufacturing build out of the build box and into a sleeve. A separation mechanism is configured to separate the portion of the additive manufacturing within the sleeve from a remainder of the additive manufacturing build. A de-powdering station includes a conveyor configured to convey the separated portion of the build away from the remaining portion of the build, and an agitation mechanism is configured to separate unbound powder build material from one or more objects of the portion while the portion is being conveyed by the conveyor.
B08B 7/02 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by distortion, beating, or vibration of the surface to be cleaned
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 64/379 - Handling of additively manufactured objects, e.g. using robots
B33Y 30/00 - Apparatus for additive manufacturingDetails thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
A recoater for an additive manufacturing apparatus includes a recoater arm and retainer operably coupled with the recoater arm. The retainer includes a housing defining a cavity. A blade carrier supports one or more blades. An actuator is operably coupled with the housing and is configured to compressively retain an upper portion of the blade carrier within the cavity between a slide of the actuator and the housing.
A gas turbine engine comprising: a fan assembly comprising a fan; a turbomachine drivingly coupled to the fan and comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine, the turbomachine defining an annular cooling passage extending between a CP inlet and a CP outlet, the CP inlet in airflow communication with the working gas flowpath, the bypass passage, or both; an accessory system; a heat exchanger positioned in thermal communication with the annular cooling passage at a location between the CP inlet and the CP outlet, the heat exchanger in thermal communication with the accessory system; and a bleed cooling system defining a BC inlet in airflow communication with the annular cooling passage at a location between the CP inlet and the CP outlet.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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
15.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, and given by
A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, and given by
V
G
V
GB
.
A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, and given by
V
G
V
GB
.
VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes a combustion section and the gearbox assembly. A fuel delivery system includes a fuel supply line for delivering fuel to the combustion section. A lubrication system includes a lubricant supply line for delivering lubricant to the gearbox assembly. A thermal management system includes a fuel-lubricant heat exchanger for cooling the lubricant with the fuel. The thermal management system selectively directs the fuel through fuel bypass lines or the lubricant through lubricant bypass lines to bypass the fuel-lubricant heat exchanger based on a fuel temperature or a lubricant temperature.
Systems, apparatus, articles of manufacture, and methods are disclosed that include a fuel distribution system for an engine, the fuel distribution system comprising: a first pump downstream of a fuel tank; a first motor coupled to the first pump; a second pump downstream of the first pump; a second motor coupled to the second pump; a third pump downstream of the second pump, the third pump having an inlet and an outlet; a bypass pathway from the outlet of the third pump to the inlet of the third pump; and a recirculation valve, the recirculation valve in line with the bypass pathway.
A turbine engine includes a fan and a turbomachine defining an engine centerline. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbine engine also includes a set of composite airfoils circumferentially arranged about the engine centerline. An airfoil in the set of composite airfoils includes a composite portion extending chordwise between a composite leading edge and a trailing edge, and a leading edge protector coupled to the composite portion.
A propulsion and electrical system for an aircraft including a turbine engine and a fuel cell. The turbine engine includes a turbo-engine, a fan having a fan shaft coupled to the turbo-engine, and a steam system. The steam system is fluidly coupled to a water source to receive water. The steam system includes a boiler to vaporize the water and to generate steam. The steam system can be fluidly coupled to the core air flow path to inject the steam into the core air flow path. The fuel cell is fluidly coupled to a hydrogen source and an oxygen source to receive hydrogen and oxygen, respectively, and, when receiving the hydrogen and the oxygen, to generate electricity and water. The fuel cell is fluidly coupled to the steam system as the water source.
An airfoil assembly for a turbine engine having a composite spar, a sleeve, and a set of lobes. The composite spar includes an exterior spar portion, where at least part of the exterior spar portion of the composite spar is received at a sleeve inner surface of the sleeve. The set of lobes extends from the composite spar and is received by a set of recesses at the sleeve inner surface.
A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. Acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the use of airfoils with leading edge protectors for both various stages of blades and vanes, the engine effectively reduces noise emissions, enhances aerodynamic efficiency, and improves structural durability.
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
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
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
22.
TURBINE ENGINE INCLUDING A FAN ASSEMBLY HAVING A DAMPER
A turbine engine experiences 1P vibrations. The turbine engine includes a turbo-engine, a compressor for compressing air, and a combustor for combusting fuel and the compressed air to generate combustion gases. A turbine receives the combustion gases and drives a fan assembly. The fan assembly includes a fan disk and a fan blade with a fan blade centerline axis. The fan blade is subjected to 1P loading, generating the 1P vibrations. A fan blade root fixed to the fan blade and connected to the fan disk, is aligned with the fan blade centerline axis, and conducts the 1P vibrations. A damper positioned radially between the fan blade root and the fan disk damps the 1P vibrations.
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 method for forming a ceramic matrix composite (CMC) article includes laying up a plurality of plies of a CMC material, sealing the plurality of plies in a vacuum enclosure that is in fluid communication with a solvent supply that includes a volume of solvent, and diffusing a solvent to or from the solvent supply through the plurality of plies.
A combustor for a gas turbine engine includes a ceramic matrix composite (CMC) dome structure, a metallic cowl structure including a single yoke outer cowl connecting flange and a single yoke inner cowl connecting flange, a CMC outer liner, and a CMC inner liner. An outer connection connects the outer cowl connecting flange, the CMC dome structure, and the CMC outer liner in an outer stacked arrangement, and an inner connection connects the inner cowl connecting flange, the CMC dome structure, and the CMC inner liner in an inner stacked arrangement.
F23R 3/60 - Support structuresAttaching or mounting means
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
An apparatus for controlling a ply of a composite material includes a robotic arm and an end effector mounted to the robotic arm. The end effector includes a base fixed to the robotic arm, a platform supported by the base, a disk supported by the platform, an arcuate track supported by the disk, and a ply manipulator extending from the arcuate track. The base includes a base track extending along a first translational dimension. The platform is movable along the base track and includes a platform track extending along a second translational dimension. The disk is movable along the platform track and is rotatable about a first rotational dimension. The arcuate track extends along a second rotational dimension. The ply manipulator is movable along the arcuate track.
An aircraft system includes an electrified powertrain having a first power module and a second power module. The aircraft system further includes a thermal management system having a thermal fluid loop for conveying a thermal fluid. The first power module and the second power module are in thermal communication with the thermal fluid loop. The thermal management system further includes a heat exchanger in thermal communication with the thermal fluid loop upstream of the first power module and the second power module such that the thermal fluid from the heat exchanger is partitioned between the first power module and the second power module.
A thermal management system defines a thermal management system flowpath to provide a flow of a fluid to an electric machine and a power electronics assembly electrically connected to the electric machine. The thermal management system includes a first heat exchanger thermally connected to the thermal management system flowpath and to the electric machine, and a second heat exchanger thermally connected to the thermal management system flowpath downstream of the first heat exchanger. The second heat exchanger is thermally connected to the power electronics assembly.
A hybrid-electric propulsion system includes a gas turbine engine comprising a high pressure system and a low pressure system, an electric machine coupled to one of the high pressure system or the low pressure system, and a thermal management system defining one or more thermal management system flowpaths to provide one or more heat exchange fluids to the electric machine. A controller collects one or more signals from one or more sensing nodes connected to at least one of the thermal management system flowpaths or the electric machine, analyzes the one or more signals to detect a thermal anomaly corresponding to at least one of the electric machine or the thermal management system, and responsive to detecting the thermal anomaly, performs at least one of adjusting a flow of at least one heat exchange fluid of the one or more heat exchange fluids or derating the electric machine.
A method of forming a coated component is provided. The method may include forming a bondcoat on a surface of a substrate; thereafter, performing a first heat treatment on the bondcoat on the surface of the substrate; thereafter, forming a barrier coating on the bondcoat; and performing a second heat treatment on the barrier coating on the bondcoat to form the coated component. The bondcoat may comprise MCrAlX where M is Ni, Co, or a combination thereof and where X is Hf, Y, Zr, or combinations thereof.
A turbine engine includes an engine core defined by a compressor section, a combustion section, and a turbine section. An inner cowl circumscribing at least a portion of the engine core is radially spaced from the engine core to define an inner cowl space. An outer cowl circumscribes at least a portion of the inner cowl. A fairing extends radially between the inner cowl and the outer cowl having at least a hollow portion. An accessory gearbox has a first portion defined by a single arm located in the inner cowl space and a second portion located in the hollow portion. A first accessory device is located in the inner cowl space and a second accessory device located in the hollow portion of the fairing.
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).
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
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
33.
PROPULSION SYSTEM USING RAMJET BLEED AIR FOR GAS TURBINE COMBUSTION
A propulsion system includes a ramjet engine and a turbine engine configured for multimode operation. The turbine engine includes a compressor section, a combustor, and a turbine section, while the ramjet engine includes a ramjet bleed air passage. A combustor inlet flow control device positioned at an outlet end of the compressor section includes a bleed valve having a ramjet bleed air inlet, a compressor bleed air outlet, and a flow control element rotatable between two positions. In a first position, the flow control device directs compressor bleed air to the turbine engine combustor. In a second position, the flow control device restricts compressor air and directs ramjet bleed air to the combustor. This configuration enables continued operation of the turbine section to drive engine accessories using ramjet-supplied air during high-speed flight when the compressor is inactive. A corresponding method for regulating airflow based on operational state is also disclosed.
A propulsion system is provided, wherein the propulsion system includes an afterburner assembly. The afterburner assembly including: an exhaust section and a fuel injector assembly that is operable to inject fuel in the exhaust section. The fuel injector assembly includes a plurality of fuel injection members. The plurality of fuel injection members defines a hot zone and a cold zone. The cold zone is positioned to provide a noise insulation barrier for the hot zone.
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluidControl thereof by after-burners
A turbine engine includes a turbo-engine, a fan including a plurality of fan blades that rotate about a longitudinal centerline axis, a rotational component coupled to the fan, a fluid circuit for supplying a fluid to the turbine engine, and a hydraulic fan brake in fluid communication with the fluid circuit. The hydraulic fan brake includes a hydraulic actuator fluidly coupled to the fluid circuit and a brake pad operably coupled to the hydraulic actuator such that the hydraulic actuator moves the brake pad, relative to the rotational component, between a disengaged position, in which the brake pad disengages the rotational component and allows rotation of the rotational component, and an engaged position, in which the brake pad engages the rotational component and prevents rotation of the rotational component, thus preventing rotation of the fan.
F04C 2/08 - Rotary-piston machines or pumps of intermeshing-engagement type, i.e. with engagement of co-operating members similar to that of toothed gearing
F15B 13/04 - Fluid distribution or supply devices characterised by their adaptation to the control of servomotors for use with a single servomotor
Oil lubricated supercritical fluid pumps with oil separators are disclosed herein. An example pump system to pressurize a fluid within a closed loop thermal transport bus disclosed herein includes a pump housing, a duct fluidly coupled to the pump housing, a first portion of the duct to include a mixture of an oil and the supercritical fluid, a second portion of the duct to include the supercritical fluid, and a separator positioned in a third portion of the duct between the first portion of the duct and the second portion of the duct, the separator to separate the oil in the mixture from the supercritical fluid.
A cooling apparatus for an electric machine includes a body, a plurality of channels defined in the body, and a manifold. The body extends from a first end to a second end and defines an outer surface, an inner surface, and a cavity interior of the inner surface. The cavity is configured to receive electric machine windings of the electric machine. The plurality of channels extend from the first end to the second end. Each of the plurality of channels is disposed between the inner surface and the outer surface. The manifold defines a fluid port and is arranged to engage the body at one of the first end or the second end to form a fluidtight chamber enclosing the plurality of channels from the cavity. The fluidtight chamber is in fluid communication with the fluid port.
H02K 3/24 - Windings characterised by the conductor shape, form or construction, e.g. with bar conductors with channels or ducts for cooling medium between the conductors
H02K 1/20 - Stationary parts of the magnetic circuit with channels or ducts for flow of cooling medium
H02K 1/32 - Rotating parts of the magnetic circuit with channels or ducts for flow of cooling medium
An extension tool including a base link comprising the proximal end, wherein the base link includes a first bend. The extension tool includes a transition section coupled to the base link and including a second bend and a plurality of sequentially arranged links coupled to the transition section and moveable relative to one another and including a third bend. The extension tool includes a support member comprising the distal end, the support member including a wheel disposed at the distal end, where the support member includes a fourth bend. The first bend of the base link extending in a first direction from a longitudinal centerline, while the second bend and the fourth bend extend in a second direction substantially along the longitudinal centerline bending downward, and the third bend of the plurality of sequentially arranged links extends in a third direction from the longitudinal centerline bending upward.
G02B 23/24 - Instruments for viewing the inside of hollow bodies, e.g. fibrescopes
F16M 13/04 - Other supports for positioning apparatus or articlesMeans for steadying hand-held apparatus or articles for supporting on, or holding steady relative to, a person, e.g. by chains
In some embodiments, a system for evaluation a coating, such as a thermal barrier coating, includes an electromagnetic inspection device and a controller in operative communication with the electromagnetic inspection device. The electromagnetic inspection device includes an electromagnetic radiation source and a detector. The electromagnetic radiation source generates pulsed electromagnetic radiation that penetrates through a coating of a component of an engine. The detector receives reflected electromagnetic radiation that is reflected from the component. The controller is configured to receive electromagnetic radiation waveform that is representative of the reflected electromagnetic radiation. The controller is also configured to determine a property of the coating based on the electromagnetic radiation waveform and to determine a remaining life of the coating based on the property. The controller may also be configured to communicate a control command to the engine based on the remaining life.
G01N 21/84 - Systems specially adapted for particular applications
G01N 21/3581 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light using far infrared lightInvestigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light using Terahertz radiation
A combustor for a gas turbine engine includes a dome structure having an outer dome connecting flange, a cowl structure, an outer liner having an outer liner connecting flange, and an inner liner having an inner liner connecting flange. At least one of (a) the outer liner connecting flange includes a bushing having an opening therethrough, or (b) the outer dome connecting flange includes a bushing having an opening therethrough. The outer dome connecting flange, the cowl structure, and the outer liner are connected together via an outer connection that includes an outer connecting member having a head and a shank. An outer surface of the head slidingly engages with at least one of (i) an inner surface of the opening of the bushing in the outer liner connecting flange, or (ii) an inner surface of the bushing in the outer dome connecting flange.
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
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
F23R 3/60 - Support structuresAttaching or mounting means
Systems, apparatus, articles of manufacture, and methods are disclosed to attach a airfoil to a hub for pitch-controlled actuation, comprising: a spar as a portion of a base of the airfoil; a trunnion coupled to the hub, the trunnion to receive the spar; the spar positioned with respect to the trunnion so that a portion of the airfoil having a length greater than a diameter of the airfoil is retained within the trunnion to react a moment through the length of the airfoil retained within the trunnion; a split ring around an inside of the trunnion, the split ring aligned with a notch on the airfoil, the notch on the airfoil shaped to allow the split ring to move towards a airfoil of the trunnion; and a nut removably coupled to the trunnion, the nut positioned to prevent the split ring from moving away from a base of the trunnion.
A textual description of an issue pertaining to at least part of an apparatus is received as input. A data store is accessed and a plurality of knowledge documents that corresponds to the input is retrieved. A language generation prompt is then generated, as a function of both the input and the plurality of knowledge documents and output to a task-specific decoder that generates a candidate recommendation to address the aforementioned issue. That candidate recommendation is output to at least one human reviewer who reviews the candidate recommendation as a function of the plurality of knowledge documents and who then provides a corresponding human-validated recommendation to address the issue. The task-specific decoder can be retrained using the human-validation recommendation coupled with the corresponding textual description input.
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.
Disclosed herein are example variable flowpath casings for blade tip clearance control. An example casing for a turbine engine includes a first annular substrate extending along an axial direction, the first annular substrate defining a cavity at a radially inward surface of the first annular substrate, a second annular substrate positioned at least partially within the cavity of the first annular substrate, and a tension belt extending circumferentially around a periphery of the second annular substrate.
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).
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
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
46.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
.
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
.
VG is a gutter volume of the gutter and VGB is a gearbox volume. The gas turbine engine includes a lubricant flow control system that includes a variable flow lubricant pump that generates a pump variable flow of lubricant to the gearbox assembly. The gearbox assembly has a variable consumption demand for delivery of lubricant. A lubricant flow controller is configured to generate a pump control command for the variable flow lubricant pump to produce the pump variable flow of lubricant based on the variable consumption demand.
F16H 57/04 - Features relating to lubrication or cooling
F16N 7/38 - Arrangements for supplying oil or unspecified lubricant from a stationary reservoir or the equivalent in or on the machine or member to be lubricated with a separate pumpCentral lubrication systems
47.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
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
G
B
.
V
G
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.
A dip-coat binder solution comprises a metal dip-coat powder and a dip-coat binder. The dip-coat binder solution has a viscosity greater than or equal to 1 cP and less than or equal to 40 cP. The metal dip-coat powder may comprise a stainless steel alloy, a nickel alloy, a copper alloy, a copper-nickel alloy, a cobalt-chrome alloy, a titanium alloy, an aluminum alloy, a tungsten alloy, or a combination thereof. A method of forming a part includes providing a green body part comprising a plurality of layers of print powder, dipping the green body part in a dip-coat binder solution to form a dip-coated green body part, and heating the dip-coated green body part. After dipping, the dip-coated green body part has a surface roughness Ra less than or equal to 10 μm.
A turbine engine includes a spool with a compressor for providing a compressed air flow to a turbine, a shaft that connects the compressor and a turbine such that the compressor and the turbine rotate together, and a fuel supply. A combustor located downstream of the compressor receives and combusts the compressed air flow and the fuel supply, generating combustion gases. The combustion gases rotate the turbine, in turn rotating the spool. A balance piston for applying an axial force on the spool includes a rotating boundary portion and a fixed enclosure portion that does not rotate, defining a balance piston cavity. A water system provides water to the balance piston cavity, heat transfer from the balance piston to the water causes expansion of the water, the water in the balance piston cavity being balance piston fluid that is pressurized and applies an axial force to the spool.
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F01D 3/00 - Machines or engines with axial-thrust balancing effected by working fluid
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 5/08 - Heating, heat-insulating, or cooling means
F02C 6/18 - Plural gas-turbine plantsCombinations of gas-turbine plants with other apparatusAdaptations of gas-turbine plants for special use using the waste heat of gas-turbine plants outside the plants themselves, e.g. gas-turbine power heat plants
F02C 7/20 - Mounting or supporting of plantAccommodating heat expansion or creep
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
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
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
51.
CATALYSTS FOR OXIDIZING COKE IN A LOW OXYGEN ENVIRONMENT
A catalyst for oxidizing coke in a low oxygen environment such as in a gas turbine engine of an aircraft. The catalyst includes a compound of formula NxM1−xO2−y. In the formula, x ranges from 0 to 0.9, y ranges from 0.02 to 0.2, N includes at least one of an alkaline-earth cation, an aluminum cation, a transition metal cation, or a rare-earth cation, M is silicon or a rare-earth element, and N has a different atomic radius than M, N has a different oxidation state than M, or N has a different atomic radius and a different oxidation state than M.
An airfoil rotor hub system including a rotor hub, the rotor hub having a variable radius around a circumference of the rotor hub, and a plurality of airfoils connected to the rotor hub, a first airfoil in the plurality of airfoils having a first span connected to a surface of the rotor hub at a first radius, and a second airfoil in the plurality of airfoils having a second span connected to the surface of the rotor hub at a second radius. The first radius is different from the second radius, and the first span of the first airfoil is different from the second span of the second airfoil so as to mistune a frequency of the first airfoil and the second airfoil to reduce or substantially to eliminate flutter in the plurality of airfoils.
A vapor generation apparatus includes a fluid channel extending between a first end and a second end of the vapor generation apparatus, a plurality of first fluid passageways extending between the first end and the second end, and a plurality of second fluid passageways extending between the first end and the second end. The plurality of first fluid passageways are between the fluid channel and the plurality of second fluid passageways. The vapor generation apparatus also includes a fluid chamber adjacent the second end and configured to receive a first fluid and a separator adjacent the first end and in fluid communication with the fluid channel, the plurality of first fluid passageways, and the plurality of second fluid passageways. The fluid chamber is in fluid communication with the fluid channel and the plurality of first fluid passageways.
Gas turbine engine including stub-tandem variable inlet guide vanes are disclosed. An example gas turbine engine includes a compressor, a variable inlet guide vane upstream of the compressor, and a part-span vane positioned between at least a portion of the variable inlet guide vane and the compressor in an axial direction defined by the gas turbine engine.
A fuel injector cooling system, comprising: a blower including a blower inlet and a blower outlet; a valve including a valve inlet and a valve outlet; a duct defining a flow passage between the blower outlet and the valve inlet; a fuel injector including a flange and a valve housing extending radially outward from the flange, wherein the valve housing defines an outer surface of the fuel injector; and a cooling jacket defining an inner surface and a cooling air inlet in fluid communication with the flow passage, wherein the cooling jacket defines a cooling flow passage between the inner surface of the cooling jacket and the outer surface of the valve housing, wherein the cooling flow passage is in fluid communication with the flow passage.
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Akkaram, Srikanth
Oracz, Dariusz Robert
Dziugiel, Tomasz
Konwar, Rajkumar Singha
Abstract
There are provided systems and methods for prognostic analytics of an asset. For example, there is provided a processor-implemented method for severity prediction for aircraft components. The method includes accessing time series flight-by-flight data relating to a component of an aircraft, the time series flight-by-flight data comprising performance data; determining, by a prediction model, an estimated degree of distress for the component based on the time series flight-by-flight data; determining a flight-by-flight severity prediction for the component based on the estimated degree of distress; and providing a preemptive recommendation for the component based on determined the flight-by-flight severity prediction.
A casting core is used in manufacture of a cast engine component. The cast engine component has a first area, a second area, a fluid passage wall separating the first area and the second area, and a connecting fluid passage extending through the fluid passage wall and interconnecting the first area and the second area. The casting core has a first core and a second core. The second core defines a first leg and a second leg of the connecting fluid passage. The second core defines a turn of the connecting fluid passage. At least a portion of the first core and an entirety of the second core is provided within a geometric boundary defined by a set of geometric characteristics of the first core and the second core within the geometric boundary.
A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the airfoil. The blade assembly has specific geometries that improve durability.
A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the airfoil. The blade assembly has specific geometries that improve durability.
An airfoil assembly for a gas turbine engine includes a first blade having a first leading edge and a first trailing edge, and a second blade circumferentially spaced from the first blade where the second blade has a second leading edge and a second trailing edge. The first blade includes a feature formed on the first leading edge where the first blade includes a chord extending from the first leading edge to the first trailing edge. The chord is located radially where a projection of a line normal to an intersection of the chord with the first trailing edge intersects the second leading edge. At least a portion of the feature is radially located at or inboard of the chord.
A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
A coupling assembly for a turbine engine. The coupling assembly includes a cold side component, a hot side component, and a fastening mechanism. The cold side component and the hot side component together at least partially form a combustion chamber. The fastening mechanism couples the hot side component to the cold side component. The fastening mechanism includes a stud disposed through the cold side component and a cap positioned on the stud. The cap defines a hollow interior and includes one or more first cap cooling holes. The one or more first cap cooling holes operably direct cooling air into the hollow interior such that the hollow interior provides a cushion of air between the combustion chamber and the stud.
A propulsion system includes a gas turbine engine includes a low speed spool and a high speed spool. The low speed spool includes a low pressure (LP) compressor coupled to a LP turbine via a LP shaft. The high speed spool comprising includes a high pressure (HP) compressor coupled to a HP turbine via a HP shaft. A first starter motor coupled to the HP shaft and configured to provide motive power to spin the HP shaft to start the gas turbine engine. A second starter motor coupled to the LP shaft and configured to provide torque to the LP shaft. A controller configured to operate on the second starter motor to provide the torque to the LP shaft based on a base LP shaft torque schedule.
An aircraft includes a combustion engine including a fuel injector, a plurality of plasma actuators disposed in the combustion engine downstream of the fuel injector, a control processing unit communicatively coupled to each plasma actuator of the plurality of plasma actuators, and at least one sensor communicatively coupled to the control processing unit, wherein the control processing unit commands the fuel injector and at least one plasma actuator of the plurality of plasma actuators to generate plasma in response to a signal from the at least one sensor.
G07C 5/08 - Registering or indicating performance data other than driving, working, idle, or waiting time, with or without registering driving, working, idle, or waiting time
General Electric Deutschland Holding GmbH (Germany)
Inventor
Chaudhari, Pushkar Chandrakant
Mn, Suma
Osama, Mohamed
Do, Kevin Michael
Abstract
A gas turbine engine having a first pressure spool and a second pressure spool includes a power transfer system to transfer mechanical power between the first pressure spool and the second pressure spool. The first pressure spool can include a first electric machine to convert mechanical power from the first pressure spool to a first electric power. A coupled electric machine having a first coupled rotor and a second coupled rotor can be rotatingly coupled to the first pressure spool and the second pressure spool, respectively. The coupled electric machine is configured to receive an output power (e.g., an output power from an AC/AC converter) to drive a winding of the coupled electric machine to enable power transfer between the first pressure spool and the second pressure spool. An AC/AC converter can be electrically disposed between the first pressure spool and the second pressure spool.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 6/00 - Plural gas-turbine plantsCombinations of gas-turbine plants with other apparatusAdaptations of gas-turbine plants for special use
H02P 7/343 - Arrangements for regulating or controlling the speed or torque of electric DC motors for regulating or controlling an individual DC dynamo-electric motor by varying field or armature current by master control with auxiliary power using Ward-Leonard arrangements in which both generator and motor fields are controlled
H02P 101/25 - Special adaptation of control arrangements for generators for combustion engines
H02P 103/20 - Controlling arrangements characterised by the type of generator of the synchronous type
67.
TURBINE ENGINE WITH A BLADE ASSEMBLY HAVING A SET OF COOLING CONDUITS
A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the airfoil. The blade assembly has specific geometries that improve durability.
A gas turbine engine having a blade assembly with a platform, an airfoil, and a shank. The airfoil has a plurality of cooling conduits, and the shank has a plurality of inlet passages to provide cooling fluid to the cooling conduits in the airfoil. The cooling fluid is vented through a plurality of cooling holes along the airfoil. The blade assembly has specific geometries that improve durability.
A turbine engine includes a fan and a turbomachine defining an engine centerline. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbine engine also includes a set of composite airfoils circumferentially arranged about the engine centerline. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
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
An engine component for a turbine engine. The engine component has a composite structure and a cover structure. The composite structure has a composite structure outer wall and a composite structure edge. The cover structure encases at least a portion of the composite structure outer wall. The cover structure has a main body. The main body extends along the at least a portion of the composite structure outer wall.
A gas turbine engine includes a compressor section and a turbine section in axial flow arrangement defining an axially extending, longitudinal centerline, and arranged as a rotor and a stator. A floating seal assembly is disposed at an interface of the rotor and the stator to seals at least portions of the rotor and the stator relative to each other. The floating seal assembly includes one or more inertial damping elements tuned to one or more frequencies.
A gas turbine engine comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising a combustor liner that at least partially defines a combustion chamber; and a gaseous fuel nozzle assembly, comprising: a steam supply to supply steam; a hydrogen fuel supply to supply gaseous hydrogen fuel; and a fuel nozzle body fluidly coupled with the steam supply and the hydrogen fuel supply.
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F02C 3/20 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
A snake-arm robot assembly may include a first link and a second link with a joint formed therebetween including first and second contact portions in rolling contact with one another to allow the first and second links to pivot with respect to one another. The first and second contact portions are configured to contact one another at first and second contact points corresponding to a first orientation and third and fourth contact points corresponding to a second orientation. The joint may be configured such that a line of action of a net force acting on the joint is incident with a first reference line extending between the first and second contact points with the links in the first orientation, and the line of action is incident with a second reference line extending between the third and fourth contact points when the links are in the second orientation.
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 fuel treatment system for a hybrid gas-electric propulsion system using a hydrocarbon fuel includes a fuel pre-treatment unit, a recuperator including a first fuel passage and a second fuel passage where the first fuel passage is in fluid communication with a fuel outlet of the fuel pre-treatment unit. A partial oxidation reformer includes a heated fuel inlet and a reformed fuel outlet. The heated fuel inlet is in fluid communication with the fuel pre-treatment unit via the first fuel passage. A solid oxide fuel cell includes an anode inlet and an anode outlet. The anode inlet is in fluid communication with the reformed fuel outlet of the partial oxidation reformer, and the anode outlet is in fluid communication with the second fuel passage of the recuperator. A combustor is in fluid communication with the anode outlet via the second fuel passage.
A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.
Methods of forming an oxide dispersion strengthened refractory-based alloy are provided. The oxide dispersion strengthened refractory-based alloy may include a refractory-based alloy comprising two or more refractory elements and forming a continuous phase; and a rare earth refractory oxide comprising at least one rare earth element and at least one of the two or more refractory elements. The rare earth refractory oxide forms discrete particles within the continuous phase, and the oxide dispersion strengthened refractory-based alloy comprises 0.1 volume % to 5 volume % of the rare earth refractory oxide.
An insertion tool is provided. The tool includes an insertion portion, a telescoping link, a joint actuation assembly, and an extension actuator. The telescoping link having a base part and an extension part, the extension part being configured to slide longitudinally relative to the base part from a retracted state to an extended state. The joint actuation assembly is configured to change an angle between the base part of the telescoping link and the insertion portion via a joint. The extension actuator is configured to actuate the extension part of the telescoping link from the retracted state to the extended state.
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.
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.
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
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
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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
82.
HEAT EXCHANGERS INCLUDING PARTIAL HEIGHT FINS HAVING AT LEAST PARTIALLY FREE TERMINAL EDGES
In an embodiment, a heat exchanger includes a monolithic body that includes a first substrate, a second substrate, a third substrate, and a plurality of partial height fins. The second substrate is arranged parallel to and spaced from the first substrate, thereby defining a first fluid flow path. The third substrate is arranged parallel to and spaced from the second substrate opposite the first substrate, thereby defining a second fluid flow path. The plurality of partial height fins extend from one of the second substrate and the third substrate toward the other of the second substrate or the third substrate, wherein a terminal edge of each partial height fin is at least partially spaced from the other of the second substrate or the third substrate.
F28F 1/26 - Tubular elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with projections, with recesses the means being only outside the tubular element and extending transversely the means being integral with the element
F28D 1/053 - Heat-exchange apparatus having stationary conduit assemblies for one heat-exchange medium only, the media being in contact with different sides of the conduit wall, in which the other heat-exchange medium is a large body of fluid, e.g. domestic or motor car radiators with the heat-exchange conduits immersed in the body of fluid with tubular conduits the conduits being straight
F28F 1/32 - Tubular elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with projections, with recesses the means being only outside the tubular element and extending transversely the means having portions engaging further tubular elements
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 measuring and marking system for a turbine engine includes a measuring apparatus and a marking apparatus. The turbine engine includes a nacelle and an abradable liner on the nacelle. The measuring apparatus includes a frame assembly having a generally radial arm portion, a generally axial arm portion, and a measuring device mounted to the generally axial arm portion to measure a profile of the abradable liner. The marking apparatus includes a mounting plate and an arm portion extending from the mounting plate. The arm portion includes a plurality of slots to indicate a plurality of known locations along an axial length of the abradable liner. The measuring apparatus is configured to measure the profile of the abradable liner through the plurality of slots in the arm portion of the marking apparatus at the plurality of known locations.
B23P 6/04 - Repairing fractures or cracked metal parts or products, e.g. castings
B23Q 17/20 - Arrangements for indicating or measuring on machine tools for indicating or measuring workpiece characteristics, e.g. contour, dimension, hardness
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G01N 29/265 - Arrangements for orientation or scanning by moving the sensor relative to a stationary material
F01D 11/14 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
F16L 55/18 - Appliances for use in repairing pipes
G01B 11/14 - Measuring arrangements characterised by the use of optical techniques for measuring distance or clearance between spaced objects or spaced apertures
B23Q 9/00 - Arrangements for supporting or guiding portable metal-working machines or apparatus
A heat exchanger for a gas turbine engine includes an outer shell extending between a first end and a second end opposite the first end. The outer shell defines a first fluid chamber, a first fluid inlet adjacent the first end, and a first fluid outlet adjacent the second end. The outer shell also defines a first plurality of second fluid inlet openings located between the first end and the second end and a second fluid outlet adjacent the second end. The heat exchanger includes a manifold disposed within the outer shell. The manifold defines a second fluid chamber and a second plurality of second fluid inlet openings. The heat exchanger defines a first fluid pathway and a second fluid pathway.
F02C 7/14 - Cooling of plants of fluids in the plant
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
A gas turbine engine includes 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 controller configured to autonomously regulate performance of the gas turbine engine in response to at least one of: a thrust demand, an energy efficiency target, or a flight profile condition.
F02C 6/06 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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 turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
L
CORE
=
[
m
(
2
0
+
m
)
*
n
(
10
+
n
)
]
(
1
1
0
0
)
*
D
CORE
+
C
I
S
.
A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
L
CORE
=
[
m
(
20
+
m
)
*
n
(
10
+
n
)
]
(
1
100
)
*
D
CORE
+
CIS
.
Methods and systems of electrochemically machining are provided. The methods may include applying a first potential to a tool electrode of an electrochemical machining system to generate a primary electric field. The electrochemical machining system may include a workpiece opposite the tool electrode, at least one bias electrode, and at least one fluid delivery channel within the at least one bias electrode. The method may further include applying at least one second potential to the at least one bias electrode. The method may further include delivering a charged electrolyte solution through the at least one fluid delivery channel into the electrolyte solution. Applying at least one second potential and the delivering the charged electrolyte solution generates at least one secondary electric field adjacent to the primary electric field and quenches at least one location of the primary electric field.
A tool stabilization mechanism is disclosed including a stabilizing device connected to a tool and movable between a deployed position and a non-deployed position, and an actuator for actuating the stabilizing device to move between the deployed position and the non-deployed position. In some forms the tool stabilization mechanism is integrated into a borescope unit, while in other forms it may be an accessory attachable to conventional borescope units. Related methods to the above are also disclosed herein.
A particle separation system includes at least one inertial particle separator spaced apart from or integrated as a unitary piece with a heat exchanger. The at least one inertial particle separator is disposed upstream of the heat exchanger. The at least one inertial particle separator is configured and arranged to direct a first fluid flow containing a first amount of particles to an inner passage provided at a hub of the heat exchanger, an outer passage provided at a tip of the heat exchanger, or both, the hub and the tip being located at a radial periphery of the heat exchanger, and to direct a second fluid flow containing a second amount of particles to a central portion of the heat exchanger to cool down a fluid circulating within the heat exchanger. The second amount of particles is substantially less than the first amount of particles.
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 7/14 - Cooling of plants of fluids in the plant
93.
LOW FRICTION COATINGS FOR BROAD TEMPERATURE RANGES
A coated component is provided that has a relatively low friction coating across a broad temperature range. The coated component includes a substrate having a surface and a wear coating over the surface of the substrate. The wear coating includes dual lubricant constituents diffused within a matrix phase. The wear coating may have an operating temperature range of 35° C. to 850° C. while having a coefficient of friction that is 0.15 to 0.5.
An intermediate investment casting mold comprising a shell having an outer surface and an inner surface. A core is located within an interior of the shell and is spaced from the inner surface to define a void between the core and the inner surface. A plurality of anchors emanate from the outer surface of the shell. The plurality of anchors include one or more of a linear anchor or a non-linear anchor.
A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
A turbomachine engine including an engine core including a high-pressure compressor, which has an exit stage having an exit stage diameter (DCORE), a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, a power turbine in flow communication with the high-pressure turbine, and a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft. The low-pressure shaft has a length (LMSR) defined by an engine core length (LCORE) given by:
L
CORE
=
[
m
(
20
+
m
)
*
n
(
10
+
n
)
]
(
1
100
)
*
D
CORE
+
CIS
.
The present disclosure is generally related to a vane assembly for an open fan engine having a rotor and a stator. The vane assembly is a plurality of vanes each arranged about the stator. Each of the vanes of the vane assembly has a leading edge (LE) with a leading edge angle (LEA). A combination of aircraft angle of attack, sideslip, and upwash due to lifting bodies can create a flow angularity into the engine. The leading edge angle (LEA) of each of the vanes varies depending upon the circumferential location about the stator so that the impact on the flow angularity into the engine is reduced or increased in different circumferential regions.
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Natarajan, Avinash
Tatiparthi, Vishnu Vardhan Venkata
Ganiger, Ravindra Shankar
G, Nagashiresha
Mathur, Prateek
Kuropatwa, Michal Tomasz
Sibbach, Arthur W.
Abstract
Systems, apparatus, articles of manufacture, and methods are disclosed to improve fan operability control using smart materials. An engine comprising an engine surface in an airflow path, a sensor positioned on the engine surface, and a smart-material-based feature positioned on the engine surface, the smart-material-based feature triggered to modify the airflow path when the sensor outputs an indication of a detected deviation from a reference value of an operating parameter of the engine.
A mixer assembly for a gas turbine engine. The mixer assembly includes a housing and a fuel injection port. The housing has a passage formed therein, and the housing includes a passage wall facing the passage. The fuel injection port is fluidly connected to a fuel source and is configured to inject a hydrocarbon fuel into the passage. At least a portion of the passage wall is a coated passage wall. The coated passage wall is (i) coated with a layer of a catalytic metal and (ii) located downstream of the fuel injection port.
A fuel mixer configured to provide a fuel-air mixture to a combustor of an engine. The fuel mixer may include a mixer body having a mixer outer wall, a center body, an annular passageway defined between the mixer outer wall and the center body, and a fuel tube assembly placed circumferentially about the mixer body. The fuel tube assembly may include at least one fuel channel for injecting a fuel flow into the annular passageway. The fuel tube assembly may be configured to cool a boundary layer flow present in the annular passageway. The fuel tube assembly may be configured to cool the mixer outer wall, the center body, or both the mixer outer wall and the center body. Heat from the mixer outer wall, the center body, or both the mixer outer wall and the center body, may pass to the fuel flow in the fuel tube assembly.
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
A system for engine effector position measurement including an effector actuator of an engine, the effector actuator comprising a house and one or more movable elements for changing an effector position of an effector coupled to the effector actuator and a time-of-flight (TOF) sensor within a housing of the effector actuator and positioned to measure a distance between the TOF sensor and a reflecting surface within the housing to determine the effector position of the effector.
