The present disclosure relates to fluid management systems and methods of calibrating the same within additive manufacturing systems. A fluid management system includes a pump and at least one fluid circuit, each fluid circuit comprising a plurality of fluid pathways, each of the plurality of fluid pathways comprising at least one flow-regulating valve and at least one actuating valve. At least a portion of each of the plurality of fluid pathways are fluidly connected by at least one actuating valve. The pump is operable to provide a fluid to each of the plurality of fluid pathways, wherein the fluid has a flowrate within each of the plurality of fluid pathways. Each flow-regulating valve is adjustable to increase or decrease the flowrate of each of the plurality of fluid pathways such that each flowrate of the plurality of fluid pathways is substantially the same.
An insertion tool is provided. The tool includes a flexible section a rigidization actuator configured to actuate the flexible section between a rigidized state and a relaxed state, wherein in the rigidized state, a fluid path is formed within the flexible section, an end effector coupled to the flexible section, and an end effector actuator comprising at least one pneumatic turbine configured to transform fluid force of a fluid from the fluid path into torque to drive a rotation of the end effector.
A turbine engine seal configured for use between a turbine engine rotor and a turbine engine static component of a turbine engine can include a seal construction having a negative thermal expansion (NTE) layer located on one or both of the turbine engine rotor and turbine engine static component. The NTE layer can include a NTE reactive component comprising a material with a negative thermal expansion coefficient. When the turbine engine rotor rubs against the turbine engine static component, heat is generated and the NTE reactive component can experience an increase in temperature from a first temperature to a second temperature. The increase in temperature causes a dimension of the NTE reactive component to decrease which consequently forms a hydrodynamic pocket useful to generate a lift force that urges separation between the turbine engine rotor and turbine engine static component. The seal construction can include a lattice compliant layer.
A composite component may include an infiltrated segment infiltrated with a molten material during a prior infiltration process, a green segment that is uninfiltrated, and a barrier segment having a microstructure different from the infiltrated segment, the green segment, or both. The microstructure of the barrier segment may be configured to slow a flow of material between the infiltrated segment and the green segment during a subsequent infiltration process.
A gas turbine engine includes a compressor section, combustion section, and turbine section is serial flow arrangement. A fuel injector supplies a mixture of fuel and air for combustion within the combustion section. An outer wall defines a mixing passage extending along a stream-wise direction including a first mixing region and a second mixing region. A first fuel passage supplies a first fuel to the first mixing region and an air passage supplies a supply of air to the first mixing region. A second fuel passage supplies a second fuel to the second mixing region.
F23D 14/36 - Burners specially adapted for use with means for pressurising the gaseous fuel or the combustion air in which the compressor and burner form a single unit
F23D 14/64 - Mixing devicesMixing tubes with injectors
F23D 14/70 - Baffles or like flow-disturbing devices
A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). The high-pressure compressor defines a high-pressure compressor exit area (AHPCExit) in square inches. 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). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.
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
A turbine engine having a compressor section, a combustor section, a turbine section, and a rotatable drive shaft. A bypass conduit couples the compressor section to the turbine section. At least one centrifugal separator is fluidly coupled to the bypass stream, where the at least one centrifugal separator includes a body, a center body, a separator inlet, and a separator outlet fluidly coupled with the turbine section to output a reduced-particle stream that is provided to the turbine section for cooling. The centrifugal separator includes an angular velocity increaser, a flow splitter, a first outlet passage defined by an inner annular wall that receives the reduced-particle stream, and an angular velocity decreaser located downstream of the flow splitter. A second outlet passage receives the concentrated-particle stream.
B01D 45/16 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by centrifugal forces generated by the winding course of the gas stream
B04C 3/00 - Apparatus in which the axial direction of the vortex remains unchanged
B04C 3/06 - Construction of inlets or outlets to the vortex chamber
F01D 5/18 - Hollow bladesHeating, heat-insulating, or cooling means on blades
F01D 9/06 - Fluid supply conduits to nozzles or the like
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 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 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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
8.
LUBRICATION SYSTEM AND METHODS OF LUBRICATING A GEARBOX ASSEMBLY
A lubrication system for a turbine engine includes a reservoir that stores a lubricant, a primary lubricant supply circuit including a primary supply pump fluidly coupled to the reservoir, and an auxiliary lubricant supply circuit including an auxiliary supply pump fluidly coupled to the reservoir. A clutch is mechanically coupled to the auxiliary supply pump, and the clutch configured to engage the auxiliary supply pump and a shaft of the turbine engine when activated. The lubrication system further includes a pressure sensor that monitors a lubricant pressure within the primary lubricant supply circuit. When the lubricant pressure within the primary lubricant supply circuit falls below a predetermined lubricant threshold, the clutch is activated to engage the auxiliary supply pump and the shaft of the turbine engine.
A fuel nozzle has a fuel nozzle body, a set of fuel jets, and a compressed air resonator. The fuel nozzle body has a central channel defining a channel centerline. The set of fuel jets extend through the fuel nozzle body. The set of fuel jets are fluidly coupled to the central channel to define an injecting section of the central channel. The compressed air resonator defines a resonating section of the compressed air channel.
A combustor for a turbine engine includes a wall defining a combustion chamber, and an acoustic damper. The acoustic damper includes a housing in fluid communication with the combustion chamber through an opening provided in the wall, the housing defining a cavity and having one or more neck holes in fluid communication with the combustion chamber, and a mechanism configured to vary a damping acoustic frequency of the acoustic damper so as to vary the damping acoustic frequency of the acoustic damper to align with an acoustic frequency of acoustic vibrations generated in the combustion chamber to attenuate an acoustic instability within the combustion chamber.
An insertion tool is provided. The tool includes a flexible section comprising a plurality of rigidizable links, an end effector actuator, and an end effector coupled to the flexible section. A flexible is shaft inserted through the flexible section, wherein torque is transferred from the end effector actuator to the distal end via the flexible shaft to cause a rotation of the end effector. A tool may include a tool-less disconnect interface between the end effector and the flexible shaft.
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, a composite structure edge, and a channel. The channel is provided along the composite structure edge. The cover encases at least a portion of the composite structure outer wall. The cover has a main body and an extension.
A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). The high-pressure compressor defines a high-pressure compressor exit area (AHPCExit) in square inches. 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). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.
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
A turbine engine includes a cooling air duct for cooling air positioned radially between a core air flow path for core air and a bypass airflow passage for bypass air. A heat exchanger is positioned in the cooling air duct to transfer heat from a heat source from within the turbine engine. The heat exchanger may be a condenser. The turbine engine may further include a steam system that extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flow path, the steam system including the condenser to transfer heat from the combustion gases to the cooling air and to condense the water from the combustion gases. The turbine engine may further include a booster fan to increase the pressure of the cooling air and the core air.
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F01D 25/32 - Collecting of condensation waterDrainage
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
A turbomachine comprises a nozzle segment including an inner shroud defining a bottom surface and a nozzle flange defining a forward side surface and an aft side surface. A floating rotor seal is coupled to the nozzle flange via a carrier flange. The carrier flange includes a forward wall and an aft wall. The nozzle flange is positioned between the forward and aft walls and a flowpath is defined therebetween. A seal pocket is defined in one of the forward wall or the aft wall and is in fluid communication with the flowpath. At least one linear seal segment is partially disposed within the seal pocket. The linear seal segment is configured to form a seal against the nozzle flange or the bottom surface in response to pressurization of the seal pocket via a working fluid in the flowpath.
Collars to retain multiple fan blades for use with an aircraft engine are disclosed herein. An example gas turbine engine includes a plurality of fan blades and a collar coupled to the plurality of fan blades. The collar includes a first mating piece and a second mating piece coupled to the first mating piece. The first and second mating pieces define an opening configured to receive a first root of a first fan blade of the plurality of fan blades, and the first and second mating pieces define a slot to retain a second root of a second fan blade of the plurality of fan blades.
A variable pitch airfoil assembly for an engine includes a disk having an annular shape extending about an axial direction, an airfoil coupled to the disk via a platform, and at least one damping element disposed between the platform and the disk. The airfoil extends outwardly from the disk in a radial direction and is rotatable relative to the disk about a pitch axis. The at least one damping element is configured to provide vibration damping by friction between the at least one damping element, the disk, and the platform while also allowing for a pitch change of the airfoil.
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
POW
=
P
pump
(
c
p
_
pump
c
p
_
water
)
(
ρ
water
ρ
pump
)
2
,
FFR
=
(
P
engine
N
H
C
fuel
)
(
c
p
_
fuel
c
p
_
pump
)
,
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
POW
=
P
pump
(
c
p
_
pump
c
p
_
water
)
(
ρ
water
ρ
pump
)
2
,
FFR
=
(
P
engine
N
H
C
fuel
)
(
c
p
_
fuel
c
p
_
pump
)
,
0.008≤POW/FFR5/3≤12, FFR is between 0.05 pounds-mass per second and 16 pounds-mass per second, and ρwater and cp_water is the density and specific heat capacity of water, respectively.
A rotor blade for a gas turbine engine is provided. The rotor blade includes a blade body formed of a first material; and a spar within a portion of the blade body, the spar formed of a second material that is different than the first material, the spar having an elongate body including a notch. The notch, weakened geometric feature, or other reduction in cross-section defines a frangible portion of the spar that is used to control a fracture of a rotor blade.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
A liner for a combustion section of a gas turbine engine includes a base portion and a stiffening portion. The base portion includes a plurality of plies of a composite material including a first ply having a fiber direction aligned with a circumferential direction of the liner and a second ply adjacent to the first ply, the second ply having a fiber direction angled away from the circumferential direction. The stiffening portion is disposed on the base portion and includes a plurality of plies of the composite material including a first ply having a fiber direction aligned with the circumferential direction, a second ply adjacent to the first ply the second ply having a fiber direction aligned with the circumferential direction, and a third ply adjacent to the second ply, the third ply having a fiber direction angled away from the circumferential direction.
A method for predicting build line locations in a part before additively manufacturing the part, includes obtaining a sliced three-dimensional model of a part for additive manufacturing, generating, for each neighboring pair of layers in the plurality of layers, a face count difference, generating, for each of the neighboring pair of layers, a surface area difference, predicting, for each of the neighboring pair of layers, that the first layer from each of the neighboring pair of layers comprises a presence of a build line based on a determination that the face count difference is less than zero and the surface area difference is greater than zero; storing a list of predicted build line layers comprising the one or more layers predicted to comprise the presence of the build line; and adjusting dimensions of the part for additive manufacturing corresponding to a layer in the list of predicted build line layers.
An insertion tool includes a housing, an elongated section at least partially within the housing, a bendable section coupled to the elongated section at, and an actuator. The actuator is configured to actuate the bendable section, via causing an axial displacement of the elongated section within the housing, from a retracted state at least partially positioned within the housing to an extended state outside of the housing. The insertion tool also includes a tensioning assembly configured to tension the bendable section into a predefined shape in the extended state. The elongated section is coupled to the housing via a rotation interface configured to cause a rotation of the bendable section during the actuation of the bendable section from the retracted state to the extended state.
A gear assembly for use with a turbomachine comprises a sun gear, a plurality of planet gears, and a ring gear. The gear assembly is connected to an input shaft and an output shaft. The sun gear is configured to rotate about a longitudinal centerline of the gear assembly, and is driven by the input shaft. A component of the gear assembly drives the output shaft. The gear assembly further comprises an output shaft reversal mechanism configured to reverse the rotational direction of the output shaft.
A variable area turbine nozzle assembly includes a guide vane including an outer centering pin defining a tab. An inner support ring is spaced radially outward from the guide vane and defines an opening and a protrusion. The protrusion is configured to engage with the tab of the outer centering pin. An outer support ring extends circumferentially around the inner support ring and defines an aperture. The outer support ring has a second coefficient of thermal expansion that is greater than or less than the first coefficient of thermal expansion. At least one linkage joins the inner support ring to the outer support ring and is configured to rotate the inner support ring circumferentially about an axial centerline of the variable area turbine nozzle assembly in response to a change in operational temperature of a combustion gas thus causing the guide vane to rotate.
Additive manufacturing apparatuses, components of additive manufacturing apparatuses, and methods of using such manufacturing apparatuses and components are disclosed. An additive manufacturing apparatus may include a recoat head for distributing build material in a build area, a print head for depositing material in the build area, one or more actuators for moving the recoat head and the print head relative to the build area, and a cleaning station for cleaning the print head.
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
B22F 10/14 - Formation of a green body by jetting of binder onto a bed of metal powder
B22F 12/00 - Apparatus or devices specially adapted for additive manufacturingAuxiliary means for additive manufacturingCombinations of additive manufacturing apparatus or devices with other processing apparatus or devices
B22F 12/17 - Auxiliary heating means to heat the build chamber or platform
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange
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
27.
APPARATUS AND SYSTEMS FOR SEPARATING PHASES IN LIQUID HYDROGEN PUMPS
Methods, apparatus, systems, and articles of manufacture are disclosed herein that include a cryogenic pump system comprising: a cryogenic liquid tank; a cryogenic pump including a suction adapter, the suction adapter connected to the cryogenic liquid tank via a liquid supply line and a gaseous return line; and a phase separator connected downstream of the cryogenic liquid tank and upstream of the cryogenic pump, the phase separator including a filtration structure integrated into the liquid supply line to separate vapor from cryogenic liquid, the phase separator connected to the gaseous return line to direct the vapor to the cryogenic liquid tank.
A coating including a plurality of indicator oxide nanoparticles, a binder, and a wetting agent. A sulfidation corrosion mitigation coating including: a sulfidation corrosion mitigation material, a binder, and a plurality of indicator oxide nanoparticles. An article including a metal alloy substrate having the sulfidation corrosion mitigation coating thereon is also provided. The sulfidation corrosion mitigation coating can include a first indicator layer containing indicator oxide nanoparticles disposed on the surface of the metal alloy substrate. Methods for inspection of an article having a coating containing a plurality of indicator oxide nanoparticles is also provided.
A gas turbine engine is provided, including: an accessory system; and a turbomachine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section collectively defining in part a working gas flowpath, the turbomachine further including: a reverse bleed system comprising a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct in fluid communication with the working gas flowpath; and an accessory cooling system including a cooling duct defining an inlet in fluid communication with the reverse bleed duct, the accessory cooling duct including a cooling tip oriented towards the accessory system to provide an airflow onto the accessory system.
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 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
Turbomachine (1) comprising an unducted propeller (14) propelling a tertiary flow (13), a fan (12) and a compressor (4) compressing a primary flow (F1), as well as an annular passage (19) for the flow of a secondary flow (F2) downstream of the fan (12); the annular passage (19) accommodating an annular row of rectifier vanes (22) and at least one heat exchanger (24) downstream of the row of vanes (22); a plurality of diffusion corridors being provided upstream of the at least one exchanger (24), each corridor being delimited circumferentially by an intrados and by an extrados of two circumferentially adjacent vanes (22), and by at least one fin carried by at least one of the two circumferentially adjacent vanes (22).
A gas turbine engine is provided. The gas turbine engine includes: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section collectively defining in part a working gas flowpath, the turbomachine further including; a reverse bleed system comprising a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct in fluid communication with the working gas flowpath; and an active clearance control (ACC) system including an inlet, a heat transfer assembly arranged around the turbine of the turbine section, and an ACC duct assembly extending from the inlet to the heat transfer assembly, the inlet of the ACC system in fluid communication with the reverse bleed duct.
A variable area turbine nozzle assembly includes a guide vane including an outer centering pin defining a tab. An inner support ring is spaced radially outward from the guide vane and defines an opening and a protrusion. The protrusion is configured to engage with the tab of the outer centering pin. An outer support ring extends circumferentially around the inner support ring and defines an aperture. The outer support ring has a second coefficient of thermal expansion that is greater than or less than the first coefficient of thermal expansion. At least one linkage joins the inner support ring to the outer support ring and is configured to rotate the inner support ring circumferentially about an axial centerline of the variable area turbine nozzle assembly in response to a change in operational temperature of a combustion gas thus causing the guide vane to rotate.
A disconnector system for disconnecting a drive shaft of a drive mechanism from rotating equipment, upon a failure of the drive mechanism or rotating equipment, includes a disconnector mechanism having a disconnector shaft disposed in a casing and moveable relative thereto, between a first position and a second position, and a cam surface on a distal end of the arm configured to engage a slidable coupler. The movement of the disconnector shaft can be triggered by an operation of a solenoid, or by a displacement of the solenoid responsive to a melting of a meltable element.
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.
F02C 3/045 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having compressor and turbine passages in a single rotor
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
35.
GAS TURBINE ENGINE HAVING A HEAT EXCHANGER LOCATED IN AN ANNULAR DUCT
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
36.
Methods and apparatus to control a surface of an aircraft engine
Methods and apparatus to control a surface of an aircraft engine are disclosed. An example system to control a surface in an aircraft engine comprises a first valve to vary a flow of cold fluid from a thermal transfer bus (TTB) to an active surface control (ASC) system based on an operating condition of the aircraft engine, the ASC system positioned adjacent to the surface, the first valve positioned upstream from the surface, and a second valve to vary a flow of hot fluid from the TTB to the ASC system based on the operating condition, the second valve positioned downstream from the surface.
A composite airfoil comprising an airfoil portion and a composite ply. The airfoil portion has an outer wall extending between a root and a tip, and between a leading edge and a trailing edge. The composite ply has a first set of fibers and a second set of fibers. Each fiber of the first set of fibers along a first centerline axis. Each fiber of the second set of fibers extends along a second centerline axis.
A gas turbine engine includes a compressor section, combustion section, and turbine section is serial flow arrangement. A fuel injector supplies a mixture of fuel and air for combustion within the combustion section. A first annular structure defines a central passage and a longitudinal axis within the fuel injector. A second annular structure is spaced from and in annular arrangement about the first annular structure to define an outer passage in annular arrangement between the first annular structure and the second annular structure.
F23D 14/36 - Burners specially adapted for use with means for pressurising the gaseous fuel or the combustion air in which the compressor and burner form a single unit
F23D 14/24 - Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other at least one of the fluids being submitted to a swirling motion
39.
LUBRICATION MAINTENANCE SYSTEMS AND METHODS OF CHANGING LUBRICATION IN A TURBINE ENGINE
A lubrication maintenance system for a gearbox assembly of a turbine engine includes a reservoir that stores a lubricant and a lubrication pump fluidly coupled to the reservoir that circulates the lubricant through the lubrication maintenance system. A heat exchanger is fluidly coupled to the lubrication pump and the gearbox assembly of the turbine engine, and a plurality of sumps are fluidly coupled to the heat exchanger. The lubrication pump is fluidly coupled to the gearbox assembly and each of the plurality of sumps, such that the lubrication pump scavenges circulated lubricant from the gearbox assembly and each of the plurality of sumps and recycles the circulated lubricant to the reservoir.
A turbine nozzle or blade includes an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber. The airfoil body has an inner surface facing the radially extending chamber. An impingement cooling structure is within the radially extending chamber. The impingement cooling structure includes: a wall spaced from the inner surface of the airfoil body; a plurality of holes defined through the wall; and a plurality of elongated thermal flex elements defined in the wall. Because the nozzle or blade is made by additive manufacturing, the airfoil body and the impingement cooling structure include a plurality of integral material layers.
A reciprocating motion insertion tool include a flexible section, an end effector actuator, a connector within the flexible section and coupled to the end effector actuator, and an end effector coupled to a distal end of the flexible section, wherein the end effector is configured to move in a reciprocating motion at a select reciprocation rate when driven by the end effector actuator.
A method of generating an engine health assessment based on one or more index values. The method includes receiving, by a first computer system, a first set of data including environmental data, and a second set of data including operating data having an operating time and location data from an aircraft having an engine, calculating localized environmental data from the first and second sets of data, generating one or more index values based on the localized environmental data, receiving, by a second computer system, the one or more index values, and a third set of data including operational data, and performing engine health assessment using an analytical model, based on the one or more index values, and the third set of data, to estimate a health of the engine, determining a future health or condition of the engine, and outputting recommended actions to perform maintenance to the engine.
43.
SYSTEMS AND METHODS FOR LIMITING VOID FORMATION IN CERAMIC MATRIX COMPOSITE COMPONENTS
A method for limiting void formation in a melt-infiltrated ceramic matrix composite (MI-CMC) component includes arranging one or more infiltrant feedstocks in fluid communication with a targeted area of the MI-CMC component. The one or more infiltrant feedstocks have a nominal melting point at or below a nominal melting point of an alloy within the MI-CMC component. The method includes heating the one or more infiltrant feedstocks to a first temperature at or above the nominal melting point of the one or more infiltrant feedstocks to form a molten phase. The method also includes infiltrating the targeted area of the MI-CMC component with the molten phase. As such, the molten phase reacts with a solid phase in the targeted area of the MI-CMC component. Further, the method includes cooling the MI-CMC component to a second temperature that is below the first temperature to solidify the molten phase.
A method of forming a preform for a composite component. The method includes laying up a plurality of plies to form an initial preform having an initial shape and partially bonding adjacent plies of the plurality of plies to each other to form a bonded preform with the initial shape. The bonded preform includes a high adherence region and a low adherence region, the bonding between adjacent plies being greater in the high-adherence region than the low-adherence region. The method also includes forming the bonded preform from the initial shape to a final shape to generate a shaped preform. The final shape has a low-contour region formed from the high adherence region and a high-contour region formed from the low-adherence region.
In one embodiment, a method includes determining a pitch value based on at least an energy goal. The energy goal may be at least one of bowed rotor motoring or an engine start. The method includes adjusting a pitch of a plurality of variable-pitch propeller blades based on the pitch value. The plurality of variable-pitch propeller blades are part of a variable-pitch propeller. The method includes, in response to a rotation of the plurality of variable-pitch propeller blades about a central axis based on the pitch of the plurality of variable-pitch propeller blades, activating an electric brake to reduce a speed of the rotation. The method further includes transferring electrical energy, generated by the electric brake reducing the speed of the rotation, to a component.
B64D 31/06 - Initiating means actuated automatically
B64D 35/04 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors
B64D 41/00 - Power installations for auxiliary purposes
A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.
An electric propulsion system includes a fan rotor assembly having a first fan portion and a second fan portion. A fan cowling defines a first flow passage and a second flow passage. An electric drive mechanism drives the fan rotor assembly. An electric drive cooling system includes a liquid coolant that provides cooling to at least a part of the electric drive mechanism, and a heat exchanger through which the liquid coolant flows is arranged within the second flow passage. The first fan portion provides a first flow of air through the first flow passage to provide at least one of a lifting force or a thrust force to the electric propulsion system, and the second fan portion provides a second flow of air through the second flow passage to provide a flow of cooling air therethrough that is in thermal communication with the heat exchanger.
A turbine engine including a turbo-engine, a gearbox assembly, a propulsor, and a lubrication system. The turbo-engine includes a compressor section, a combustor, a turbine section, and an input shaft. The gearbox assembly includes a first gear, a plurality of second gears, and a third gear. The propulsor has an output shaft drivingly coupled to the input shaft through the gear assembly. The lubrication system is characterized by a Gearbox Lubrication System Parameter (GLSP) between 0.2 and 140 when a mass flow rate of the lubricant is linear with a lubricant pump speed, where the GLSP is given by:
The GLSP is between 0.2 and 70 when the mass flow rate of the lubricant is modulated, where the GLSP is given by:
.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 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
A de-powdering system includes one or more sidewalls defining a support chamber configured to contain an additive manufacturing build where the additive manufacturing build includes one or more objects disposed within a powder build material. A fluidization mechanism is fluidically couplable to a fluid source and includes one or more flow channels fluidically coupled to the support chamber. The fluid source is actuatable to provide a fluid from the fluid source to the support chamber and inject the fluid into the support chamber via the one or more flow channels. The one or more flow channels are oriented to introduce a swirling flow of the fluid into the support chamber to fluidize at least a portion of the powder build material within the support chamber.
An additive manufacturing apparatus includes a stage configured to hold a component. A radiant energy is device operable to generate and project radiant energy in a patterned image. An actuator is configured to change a position of the stage relative to the radiant energy device. A deposition assembly is upstream of the stage and configured to deposit a resin on a resin support. The deposition assembly includes a reservoir housing configured to retain a volume of resin between the upstream wall and the downstream wall. The deposition assembly also includes an application device operably coupled with the reservoir housing. A computing system is operably coupled with the application device. The computing system is configured to intermittently initiate a flush operation between successive layers of the component, wherein the application device is moved from a first position to a second position during the flush operation.
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
A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes 18-26 fan blades. The low-pressure turbine includes four rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-5.1. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.2-1.3.
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 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
A method of manufacturing a composite panel is provided. The method includes applying a composite face sheet to a first side of a core structure, the core structure comprising a plurality of first ceramic particles each having a first particle size that is within a first particle size range and the composite face sheet comprising a plurality of second ceramic particles each having a second particle size that is within a second particle size range, wherein the second particle size range is smaller than the first particle size range and densifying the composite panel through infiltration, wherein the infiltration comprises transport of an infiltrant through the core structure and into the composite face sheet.
An airfoil includes a non-uniform weave structure that is two-dimensional or three-dimensional, the non-uniform weave structure including a plurality of reinforcing fibers, the non-uniform weave structure having a first region with a first stiffness and a second region with a second stiffness higher than the first stiffness, wherein the plurality of reinforcing fibers include higher density fibers in the second region and lower density fibers in the first region so as to increase a stiffness of the airfoil at the second region of the airfoil in a desired orientation to achieve a desired aeromechanics response of the airfoil.
A powder removal apparatus includes an extraction housing comprising a sidewall that is sized and configured to extend around a powder bed of a build module and a top wall that is sized and configured to extend between opposite sides of the sidewall and over the powder bed. The sidewall and top wall are configured to form a chamber portion of a turbulence chamber. The top wall has a vacuum exit opening that is configured to fluidly connect to a vacuum source. The sidewall has a plurality of sidewall inlet flow channels that extend from an inlet opening at an exterior side of the sidewall to an outlet opening at an interior side of the sidewall. A side exit channel is configured to extend along the top wall from a collector opening in communication with the chamber portion toward the vacuum exit opening.
B22F 12/00 - Apparatus or devices specially adapted for additive manufacturingAuxiliary means for additive manufacturingCombinations of additive manufacturing apparatus or devices with other processing apparatus or devices
B29C 64/255 - Enclosures for the building material, e.g. powder containers
B29C 71/00 - After-treatment of articles without altering their shapeApparatus therefor
B33Y 30/00 - Apparatus for additive manufacturingDetails thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
55.
BUILD MATERIAL ESCAPEMENT ASSEMBLY AND ADDITIVE MANUFACTURING SYSTEMS INCLUDING SAME
A build material escapement assembly for an additive manufacturing system includes a retaining plate defining an outer perimeter and an inner perimeter, a retractable plate, and a top plate coupled to the retractable plate, the top plate including a top plate perimeter and being actuatable between a retracted position and an extended position. A diaphragm is coupled to the top plate via the retractable plate and further includes an exposed area extending between the inner perimeter of the retaining plate and the top plate perimeter of the top plate. The retractable plate is actuated in a lateral direction as the top plate is moved from the retracted position to the extended position, and a width of the exposed area of the diaphragm is greater than a travel distance that the retractable plate is actuated in the lateral direction.
B29C 64/307 - Handling of material to be used in additive manufacturing
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
An additive manufacturing system for additively manufacturing a three-dimensional object includes a beam generation device and a first lens array disposed downstream from the beam generation device. The first lens array divides an energy beam received from the beam generation device into a plurality of beam segments. An optical modulator is disposed downstream from the first lens array and is modulated to reflect or transmit one or more beamlets from the plurality of beam segments incident on the optical modulator. A second lens array is disposed downstream from the optical modulator where the one or more beamlets are incident upon the second lens array. A focusing lens assembly is disposed downstream from the second lens array. The one or more beamlets projected from the second lens array become incident on the focusing lens assembly, and the focusing lens assembly converges the one or more beamlets in a target plane.
B29C 64/268 - Arrangements for irradiation using laser beamsArrangements for irradiation using electron beams [EB]
B33Y 30/00 - Apparatus for additive manufacturingDetails thereof or accessories therefor
G02F 1/01 - Devices or arrangements for the control of the intensity, colour, phase, polarisation or direction of light arriving from an independent light source, e.g. switching, gating or modulatingNon-linear optics for the control of the intensity, phase, polarisation or colour
57.
AUTOMATED DE-POWDERING OF ADDITIVE MANUFACTURING BUILD
A de-powdering system for additively manufactured objects includes an enclosure defining a cavity configured to support an additive manufacturing build. The enclosure includes a wall defining a lower boundary of the cavity, and the wall includes one or more flow channels. A sleeve is disposable in the cavity to at least partially surround the additive manufacturing build. At least one vibration mechanism is coupled to the sleeve and is actuatable to induce vibrations to the additive manufacturing build to loosen at least a portion of a powder build material from one or more objects suspended within the powder build material. The powder build material is removed from the cavity via the one or more flow channels.
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
B08B 13/00 - Accessories or details of general applicability for machines or apparatus for cleaning
B08B 15/02 - Preventing escape of dirt or fumes from the area where they are producedCollecting or removing dirt or fumes from that area using chambers or hoods covering the area
An apparatus and method for an inspection apparatus for inspecting an engine component. The inspection apparatus includes at least one controller configured to receive a set of inspection parameters based on a detection metric. A non-destructive evaluation (NDE) instrument for scanning a predetermined area of a surface of the engine component according to the set of inspection parameters to generate a data set is included. Further, the inspection apparatus includes a computer configured to apply a detection algorithm to the data set.
G01N 23/223 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by measuring secondary emission from the material by irradiating the sample with X-rays or gamma-rays and by measuring X-ray fluorescence
A hybrid electric gas turbine engine is provided. The hybrid electric gas turbine engine includes: a turbomachine having a compressor section and a turbine section arranged in serial flow order, the compressor section and turbine section together defining a core air flowpath, the turbomachine defining a core air flowpath exhaust; and an electric machine assembly having an electric machine disposed aft of the core air flowpath exhaust and mechanically connected to the turbine section.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
A 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). The gas turbine engine further includes a blade effective acoustic length (BEAL), an acoustic spacing, and an acoustic spacing ratio (ASR). The ASR can be in a range from 1.5 to 16.0.
A method of mitigating rotor bow in a rotor of a turbine engine. The method includes determining thermal rotor bow in the rotor, determining non-thermal rotor unbalance in the rotor by monitoring the response at the bowed rotor mode of the rotor, and determining a time period for motoring the rotor prior to operation of turbine engine, wherein the time period is based on a combination of the thermal rotor bow and the non-thermal rotor unbalance. The method also includes motoring the rotor for the time period and until the vibration of the rotor is below a predetermined acceptable value.
F01D 19/02 - Starting of machines or enginesRegulating, controlling, or safety means in connection therewith dependent on temperature of component parts, e.g. of turbine casing
A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section. The turbine section includes a band having an upstream end and a downstream end. The band extends between the upstream end and the downstream end, and the band at least partially defines the working gas flow path. A plurality of airfoils extend into the working gas flow path from the band. Each airfoil of the plurality of airfoils includes a leading edge, a trailing edge, a first side, and a second side opposite the first side. Each of the plurality of airfoils is substantially symmetric across an airfoil centerline extending through a center of each of the plurality of airfoils. The band defines a valley portion adjacent the leading edge of each of the plurality of airfoils and a pair of hill portions on opposing sides of the valley portion.
A turbine engine including a fan having a plurality of fan blades, a turbo-engine positioned downstream of the fan and having a core inlet, an engine intake that extends to the core inlet, and a variable engine intake system. Air enters the turbine engine through the engine intake. The variable engine intake system includes a plurality of wind condition sensors for sensing wind conditions about the turbine engine, and a controller. The plurality of wind condition sensors includes a first wind condition sensor on a first side of the turbine engine and a second wind condition sensor on a second side of the turbine engine. The controller adjusts the engine intake based on the wind conditions about the turbine engine from the first wind condition sensor and the second wind condition sensor.
Systems, apparatus, articles of manufacture, and methods are disclosed that include an air foil bearing, the air foil bearing comprising: a thrust disc coupled to a rotor shaft, the thrust disc and rotor shaft to rotate; a thrust pad aligned with a first side of the thrust disc, the thrust pad to engage with the thrust disc as the thrust disc rotates; and a micro lattice structure between the thrust disc and the thrust pad, the micro lattice structure to mitigate the thrust pad engaging with the thrust disc.
A combustor of a turbine engine includes a first combustion zone operable to combust a first fuel and air mixture, a first fuel inlet for providing a first fuel, a first air inlet for providing first zone air, the first fuel and the first zone air combining to form the first fuel and air mixture in the first combustion zone, a second combustion zone operable depending on turbine engine operating parameters, for combusting a second fuel and air mixture, a second fuel inlet providing a second fuel, that operates when the second combustion zone is operating and does not operate when the second combustion zone is not operating, and a second air inlet providing second zone air, the second fuel and the second zone air combining to form the second fuel and air mixture in the second combustion zone, wherein the first fuel and the second fuel are disparate fuels.
A turbine engine operable in a cold start condition to prevent wear on components of a gearbox assembly of the turbine engine. The turbine engine includes a gearbox assembly, a pump for directing lubricant to the gearbox assembly, a supply line heating path comprising a heat exchanger, a recirculation bypass path, a valve being positionable between a first position to direct the flow of the lubricant into the supply line heating path and a second position to direct the flow of the lubricant into the recirculation bypass path, and an electronic control unit configured to position the valve between the first position and the second position based on a temperature of the lubricant.
Variable displacement pumps and related methods are disclosed herein. An example pump case disclosed herein, the pump case defining a fluid pathway between an inlet and an outlet, the pump case including an interior surface, a shaft, and a vane disposed adjacent to interior surface, the vane coupled to the shaft, the vane including a metallic core, and a ceramic interface coupled to at least one of a surface of the metallic core, a tip of the vane, or the interior surface.
F04C 2/344 - Rotary-piston machines or pumps having the characteristics covered by two or more of groups , , , or having the characteristics covered by one of these groups together with some other type of movement between co-operating members having the movement defined in group or and relative reciprocation between the co-operating members with vanes reciprocating with respect to the inner member
68.
THERMAL MANAGEMENT SYSTEM FOR A GAS TURBINE ENGINE
A thermal management system for a gas turbine engine includes a heat exchanger including first and second sides, with the first side in contact with flow path air flowing through a flow path of the engine. Furthermore, the system includes a housing positioned relative to the heat exchanger such that the housing and the second side of the heat exchanger define a plenum configured to receive bleed air from the engine. Moreover, the system includes and at least one of a plurality of fins extending outward from the second side of the heat exchanger in a radial direction into the plenum and along the second surface of the heat exchanger in the circumferential direction or an impingement plate defining a plurality of impingement apertures, with each impingement aperture configured to direct an impingement jet of the bleed air within the plenum onto the second side of the heat exchanger.
A bulk dual phase soft magnetic component having a three-dimensional magnetic flux and its manufacturing methods are described herein. The methods can include combining a first powder material with a second powder material to form a component structure, wherein the first powder material comprises a plurality of first particles each comprising a first core and a reactive coating, and wherein the second powder material comprises a plurality of second particles each comprising a second core and a non-reactive coating, and, consolidating the component structure to join the plurality of first particles with the plurality of second particles.
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
B22F 1/16 - Metallic particles coated with a non-metal
H01F 1/00 - Magnets or magnetic bodies characterised by the magnetic materials thereforSelection of materials for their magnetic properties
H01F 41/02 - 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
A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of a turbine to receive the combustion gases and to condense water. The fuel heat exchanger is thermally coupled to the condenser to receive heat from the water condensed by the condenser. The fuel heat exchanger is located in the fuel delivery assembly to receive the fuel and to transfer the heat received from the water to the fuel. The boiler is located downstream of the fuel heat exchanger. The boiler receives the water and is fluidly connected to the combustor to receive the combustion gases and to boil the water to generate steam.
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
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
F01K 23/10 - Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids the engine cycles being thermally coupled combustion heat from one cycle heating the fluid in another cycle with exhaust fluid of one cycle heating the fluid in another cycle
A turbine engine includes a turbomachine having a compressor section, a combustor, and a turbine section. The turbine engine includes a set of composite airfoils, where a composite airfoil of the set of composite airfoils includes a composite body that extends chordwise between a composite leading edge and a trailing edge. A leading edge protector is coupled to the composite body. A platform extends from the composite airfoil.
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange
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
A composite airfoil for a turbine engine, the composite airfoil having at least one airfoil body element. The at least one airfoil body element includes a core and a composite wrap, where the composite wrap overlies at least a portion of the core. The core includes a set of composite plies.
General Electric Company Polska sp. z o.o. (Poland)
Inventor
Sibbach, Arthur W.
Łobocki, 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.
Methods and apparatus for anti-ice heat supply from waste heat recovery systems are disclosed. An apparatus for an aircraft, the apparatus comprising a fuel heat exchange system, an anti-ice heat exchanger, a waste heat recovery heat exchanger, and a conduit coupled to the fuel heat exchange system, the anti-ice heat exchanger, and the waste heat recovery heat exchanger, the conduit including a first portion and a second portion distinct from the first portion, wherein the first portion of the conduit carries a first portion of a thermal transfer fluid from the waste heat recovery heat exchanger to the anti-ice heat exchanger in which the thermal transfer fluid supplies anti-ice heat to a portion of the aircraft, wherein the second portion of the conduit carries a second portion of the thermal transfer fluid from the waste heat recovery heat exchanger to the fuel heat exchange system.
A casing for a turbine engine including a composite section. The composite section has (i) an arcuate shape or an annular shape and (ii) a circumferential direction. The composite section includes a matrix and a plurality of circumferential reinforcing fiber tows embedded in the matrix. Each circumferential reinforcing fiber tow of the plurality of circumferential reinforcing fiber tows extends in the circumferential direction and has a plurality of undulations in the circumferential direction to allow the matrix material of the composite section to expand circumferentially.
A turbine engine including a fan having a plurality of fan blades, a nacelle that extends circumferentially about the fan, an engine intake including an engine inlet, and a variable engine intake system. The nacelle includes a fan cowl and an inlet cowl that is movable with respect to the fan cowl. The engine inlet is defined from a leading edge of the nacelle to the plurality of fan blades. The variable engine intake system adjusts the inlet cowl axially between a fully retracted position and a fully extended position to adjust an inlet length of the engine inlet. The inlet cowl maintains contact with the fan cowl when the inlet cowl is extended.
A 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 HPC defining a HPC exit area (AHPCExit) in square inches, the turbine section having a LPT; a fan; a gearbox; and a DECP 0.10-0.50, where DECP equals 10*D/GR/NLPT2/NHPC, D is the fan blade tip diameter, GR is the gear ratio of the gearbox, NLPT is the stage count of the LPT, and NHPC is the stage count of the HPC. 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).
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 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
79.
AUTOMATED FIBER PLACEMENT ASSEMBLY WITH PRESSURE ROLLER
The disclosure herein relates to an automated fiber placement assembly for forming a component by the successive layering of strips of fiber tows with a pressure roller. The pressure roller has a rotational axis about which the pressure roller rotates to apply the strip of fiber tows to the component. The pressure roller can be shaped complementary to a non-uniform surface of the component for even application of the strip of fiber tows to the component.
B29C 70/38 - Automated lay-up, e.g. using robots, laying filaments according to predetermined patterns
B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or coreShaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
Structures for achieving reduced span of tie rods and improved vibration mode margins in gas turbine engines are described. The gas turbine engine includes a tie rod assembly, a plurality of coupling nuts, a forward shaft, a blisk, a thread engagement coupled to a cone shaft of the blisk, a high pressure compressor rotor, and a high pressure turbine rotor comprising a cone shaft. A first coupling nut is coupled to the cone shaft of the high pressure compressor rotor. A second coupling nut is coupled to the forward shaft. A third coupling nut is coupled to an aft end stage of the high pressure turbine rotor.
Structures for achieving high bypass ratio in gas turbine engines are described. A gas turbine engine includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section includes a first booster, a second booster, and a blisk including an inclined web and offset bore. An angle of inclination relative to a vertical plane is between 5° and 55°. The vertical plane is perpendicular to an axial direction that is parallel to a longitudinal centerline defined by the gas turbine engine. The turbine section includes a first turbine and a second turbine.
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
A composite panel assembly includes a first composite panel having a first core structure. The first core structure includes a first core structure first face and a first core structure second face. The first composite panel further includes a first composite panel first composite sheet and a first composite panel second composite sheet. The first composite panel first composite sheet bonded to the first core structure first face and a first composite panel second composite sheet bonded to the first core structure second face. An interlocking feature is defined in the first core structure. The composite panel assembly further includes a component having a portion that extends into the interlocking feature such that an interlocking mechanical joint is formed between the first composite panel and the component.
B32B 3/06 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by features of form at particular places, e.g. in edge regions for securing layers togetherLayered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by features of form at particular places, e.g. in edge regions for attaching the product to another member, e.g. to a support
B32B 3/12 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. apertured or formed of separate pieces of material characterised by a layer of regularly-arranged cells whether integral or formed individually or by conjunction of separate strips, e.g. honeycomb structure
B32B 9/00 - Layered products essentially comprising a particular substance not covered by groups
B32B 9/04 - Layered products essentially comprising a particular substance not covered by groups comprising such substance as the main or only constituent of a layer, next to another layer of a specific substance
B33Y 80/00 - Products made by additive manufacturing
83.
COMPOSITE PANELS HAVING AN INTEGRATED ATTACHMENT FEATURE AND METHODS FOR MAKING THE SAME
A composite panel includes a core structure having a main body and an attachment portion integrally formed with the main body. The core structure further includes at least one face. The attachment portion defines a first portion of an attachment aperture. The composite panel further includes a composite sheet that is bonded to the least one face of the core structure. The composite sheet extends between the main body and the attachment portion. The composite sheet defines a second portion of the attachment aperture.
B32B 3/04 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by features of form at particular places, e.g. in edge regions characterised by a layer folded at the edge, e.g. over another layer
B32B 3/12 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. apertured or formed of separate pieces of material characterised by a layer of regularly-arranged cells whether integral or formed individually or by conjunction of separate strips, e.g. honeycomb structure
B32B 3/26 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a particular shape of the outline of the cross-section of a continuous layerLayered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a layer with cavities or internal voids
B32B 3/30 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a particular shape of the outline of the cross-section of a continuous layerLayered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a layer with cavities or internal voids characterised by a layer formed with recesses or projections, e.g. grooved, ribbed
B32B 9/00 - Layered products essentially comprising a particular substance not covered by groups
B32B 9/04 - Layered products essentially comprising a particular substance not covered by groups comprising such substance as the main or only constituent of a layer, next to another layer of a specific substance
B33Y 80/00 - Products made by additive manufacturing
84.
COMPOSITE TUBE ASSEMBLIES AND METHOD OF MANUFACTURING
A composite tube assembly includes a first composite tube having a first tubular core and a first outer composite material. The first outer composite material is bonded to a first outer face of the first tubular core. The composite tube assembly further includes a second composite tube having a second tubular core and a second outer composite material. The second outer composite material is bonded to a second outer face of the second tubular core. The second composite tube is coupled to the first composite tube.
A gas turbine engine includes a fan assembly, a turbomachine defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine, a core cowl, and a heat exchanger assembly including a heat exchanger and a heat exchanger cowl defining a cooling air flowpath extending between a flowpath inlet in airflow communication with the bypass passage to receive a cooling airflow from the bypass passage and a flowpath outlet in airflow communication with the bypass passage to exhaust the cooling airflow back to the bypass passage, the heat exchanger positioned within the cooling air flowpath, the cooling air flowpath comprising a diffusion section located between the flowpath inlet and the heat exchanger.
An aeronautical vehicle comprising: a vehicle body; a propulsion system operable with the vehicle body, the propulsion system comprising a gas turbine engine, the gas turbine engine comprising a combustion section; a fuel delivery assembly comprising a gaseous fuel delivery section extending to the combustion section, the gaseous fuel delivery section defining a potential leak location; and a fuel leak combustion assembly comprising a heat source positioned in communication with potential leak location to ignite a gaseous fuel leaking from the potential leak location of the gaseous fuel delivery section of the fuel delivery assembly.
A turbine engine for an aircraft includes a fuel delivery assembly for a hydrocarbon fuel to flow therethrough, a combustor combusting the fuel to generate combustion gases, and a core air exhaust nozzle exhausting the combustion gases from the turbine engine. The turbine engine also includes a contrail mitigation system having a heater and a fuel precipitate separator. The heater is selectively operable to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel, and the fuel precipitate separator separates the fuel precipitates generated by the heater from the fuel. A controller is coupled to the heater to operate the heater to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel in response to a contrail mitigation input.
A gas turbine engine comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a gaseous fuel nozzle assembly, comprising: a rich fuel supply to supply a rich mixture of gaseous fuel and air; a lean fuel supply to supply a lean mixture of gaseous fuel and air; a rich impingement tube fluidly coupled to the rich fuel supply and emitting the rich mixture into the combustion chamber; and a lean impingement tube fluidly coupled to the lean fuel supply and emitting the lean mixture into the combustion chamber.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
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
F23D 14/58 - Nozzles characterised by the shape or arrangement of the outlet or outlets from the nozzle, e.g. of annular configuration
89.
POWER CONVERTER AND SYSTEM FOR AN ENGINE STARTER GENERATOR
A power converter system includes an asynchronous induction generator electrically coupled to an inverter/converter/controller (ICC). The ICC is coupleable to an electrical load. The ICC can include an AC-DC converter and a DC-DC converter. The DC-DC converter is configured to operate at a duty cycle substantially equal to 1 in a first operating mode. In the event of a short-circuit fault in the electrical load, the DC-DC converter is configured to operate at a duty cycle less than 1 in a second operating mode.
H02M 7/04 - Conversion of AC power input into DC power output without possibility of reversal by static converters
H02M 3/155 - Conversion of DC power input into DC power output without intermediate conversion into AC 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
90.
SUB-COOLERS FOR REFUELING ONBOARD CRYOGENIC FUEL TANKS AND METHODS FOR OPERATING THE SAME
A sub-cooler for a sub-cooling cryogenic refueling system is disclosed herein. An example method to refuel an onboard cryogenic fuel tank by controlling a sub-cooler of a cryogenic refueling system, the method comprising determining, using a first controller, a commanded first valve actuator position based on at least a source temperature and a target temperature, determining, using the first controller, an error between a measured temperature from a temperature sensor and the target temperature, determining, using the first controller, the commanded first valve actuator position based on the error and a preceding commanded first valve actuator position, determining, using a second controller, an actual first valve actuator position based on the commanded first valve actuator position, and generating, using the second controller, a primary first valve effective area and an auxiliary first valve effective area based on the actual first valve actuator position.
Methods and systems of electrochemically machining a component are provided. The method may include applying two or more potentials to a tool electrode comprising an array of two or more individual electrodes to generate two or more electric fields in between the tool electrode and a workpiece opposite of the tool electrode, wherein each of the two or more electric fields is generated by one of the array of two or more individual electrodes.
A gas turbine engine (10) defining a radial direction and an axial direction, the gas turbine engine (10) comprising: a fan assembly (140) comprising a fan; a turbomachine (16) 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 (37), the gas turbine engine (10) defining a bypass passage (56) over the turbomachine (16); a core cowl (112); and a heat exchanger assembly (124) disposed aft of the fan in the axial direction and forward of the core cowl (112) in the axial direction, the heat exchanger assembly (124) comprising: a heat exchanger (126); and a heat exchanger cowl (138) defining a cooling airflow flowpath extending between a flowpath inlet (108) in airflow communication with the bypass passage (56) to receive a cooling airflow from the bypass passage (56).
A turbine airfoil comprises an airfoil defining a leading edge, a trailing edge, a root portion, a tip portion, a chord line defining a chord length of the airfoil, a suction side surface extending in a spanwise direction from the root portion to the tip portion and in a flow-wise direction between the leading edge and the trailing edge, and a throat line extending spanwise along the suction side surface from the root portion to the tip portion. The turbine airfoil further includes a plurality of micro-riblet patches defined along the suction side surface aft of the throat line where each micro-riblet patch of the plurality of micro-riblet patches extends in the flow-wise direction between the throat line and the trailing edge.
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. Relationships between acoustic spacing and a high-speed shaft rating allow for a gas turbine engine that reduces noise emissions while maintaining high performance.
F04D 29/66 - Combating cavitation, whirls, noise, vibration, or the likeBalancing
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
95.
Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes
A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The combination of acoustic spacing and corrected specific thrust provide enhanced propulsive efficiency.
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 seal assembly defines a primary seal at an interface of the rotor and the stator and separates an inlet plenum from an outlet plenum. The seal assembly includes one or more aspiration conduits fluidly connected to the primary seal. One or more flingers are positioned at least partially between the seal assembly and the inlet plenum that define one or more passageways fluidly connecting the inlet plenum to the one or more aspiration conduits. The flingers are positioned to direct at least a portion of an airflow within the inlet plenum away from the one or more passageways.
A lubrication system for a turbine engine that includes one or more rotating components. The lubrication system includes one or more tanks that store lubricant, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during stable operating conditions of the lubrication system. The auxiliary lubrication system includes an auxiliary feed line and an auxiliary supply line. The auxiliary lubrication system receives the lubricant from the one or more tanks through the auxiliary feed line. The auxiliary lubrication system supplies the lubricant to the one or more rotating components through the auxiliary supply line when there is a potential lubricant interruption in the lubrication system.
Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 7/10 - Heating air supply before combustion, e.g. by exhaust gases by means of regenerative heat-exchangers
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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
99.
Mixing elements for rotating detonation combustion systems
A rotating detonation combustion system includes a detonation channel including an inner wall and an outer wall and extend in a longitudinal direction from an inlet of the detonation channel to an outlet of the detonation channel. A first mixing element and a second mixing element are disposed in the rotating detonation combustion system. The first mixing element forming a ring on the inner wall and the second mixing element forming a ring on the outer wall of the detonation channel adjacent the inlet. Each of the first mixing element and the second mixing element comprise a plurality of protrusions disposed circumferentially along the inner wall and the outer wall and extend into the detonation channel such that the plurality of protrusions affects vectors of at least a portion of the fuel and at least a portion of the fluid passing through recesses between the plurality of protrusions.
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
