A method of operation is provided during which a fluid is directed through a fluid line servicing a component of an aircraft engine. Pressure wave oscillations travel through the fluid within the fluid line at a frequency equal to or greater than one thousand five hundred hertz during the directing of the fluid. The pressure wave oscillations traveling through the fluid within the fluid line are damped using an elastomeric foam sleeve in contact with the fluid line. The elastomeric foam sleeve extends longitudinally along and circumscribes the fluid line.
A method of forming z-channels in a fibrous ceramic preform includes mounting the preform in a tooling assembly, the tooling assembly comprising a first fixture and a second fixture, heating the preform mounted in the tooling assembly to a temperature above a glass transition temperature of a polymer binder within the preform to induce a softened state of the polymer binder, inserting a plurality of needles sequentially through respective first holes in the first fixture, the preform, and respective corresponding second holes in the second fixture, and removing the plurality of needles leaving behind a corresponding plurality of z-channels in the preform.
A tooling assembly for use in forming z-channels in a fibrous ceramic preform includes a mandrel having a first plurality of holes extending into a mandrel body, a first subset of the plurality of holes being through-holes extending completely through the mandrel, and a second subset of the holes comprising blind pockets, and a plurality of channels extending longitudinally along the mandrel in a direction orthogonal to the first plurality of holes. The tooling assembly further includes an outer fixture at least partially enclosing the mandrel, the outer fixture including at least one piece comprising a second plurality of holes extending completely through the at least one piece, the second plurality of holes being aligned with respective corresponding ones of the first plurality of holes such that a needle can be inserted through each of the second plurality of holes in the at least one piece and into the respective ones of the first plurality of holes in the mandrel.
A method for non-destructive testing and measurement of corrosion attacks includes defining characteristic corrosion attack parameters, exposing a first specimen to corrosive conditions to induce multiple corrosion attack sites, measuring the time of exposure to corrosive conditions, measuring one or more spatially resolved corrosion attack characteristic parameters for the multiple corrosion attack sites to provide a first corrosion data set. The first set of spatially resolved corrosion attack characteristic parameters are measured by a non-destructive technique and the probability of failure for the first specimen from the first corrosion data set is modeled. The composition of the specimen may be changed based on results achieved.
G06F 119/02 - Reliability analysis or reliability optimisationFailure analysis, e.g. worst case scenario performance, failure mode and effects analysis [FMEA]
5.
LOCATION-SPECIFIC PROBABILISTIC APPROACH FOR PREDICTION OF DEFECT FORMATION IN ADDITIVE MANUFACTURING
A method for location-specific probabilistic prediction of formation of a defect in powder bed fusion additive manufacturing of an article includes determining first statistical distributions of multiple process parameters selected from laser power, scan speed, laser spot size, and powder layer thickness and density, based on the first statistical distributions, for locations across the article, determining second statistical distributions of a threshold temperature (Tthresh) for formation of the defect at each of the locations, determining a cumulative temperature thermal history (T0) at each of the locations across the article, for each of the locations across the article, determining a probability of formation of the defect based upon a probability of Tthresh versus T0, and from the probability of formation of the defect at each of the locations, generating an article integrity map.
In a method for manufacturing a turbine engine element such as a blade or vane, the element has an airfoil. The method includes: applying a load across an assembly of a first cast portion of the airfoil and a second cast portion of the airfoil; and applying current across a junction of the first cast portion and the second cast portion to fuse the second cast portion to the first cast portion.
B22D 19/16 - Casting in, on, or around, objects which form part of the product for making compound objects cast of two or more different metals, e.g. for making rolls for rolling mills
B22D 21/00 - Casting non-ferrous metals or metallic compounds so far as their metallurgical properties are of importance for the casting procedureSelection of compositions therefor
B23K 9/095 - Monitoring or automatic control of welding parameters
B23K 11/00 - Resistance weldingSevering by resistance heating
B23K 13/01 - Welding by high-frequency current heating by induction heating
B23K 101/00 - Articles made by soldering, welding or cutting
B23P 15/04 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
7.
PRE-SHAPED SPACE FILLERS FOR SMALL CROSS SECTION FEATURES
A method of fabricating a variable geometry space filling insert for a ceramic matrix composite component includes arranging a plurality of fiber bodies into a preform insert, applying a polymer binder to the plurality of fiber bodies, trimming at least a first subset of the plurality of fiber bodies such that each of the first subset of fiber bodies is shorter than at least a second subset of fiber bodies, shaping the insert with a forming tool to form a shaped insert.
C04B 35/657 - Processes involving a melting step for manufacturing refractories
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
8.
COMPONENT CORROSION PROGNOSTICS USING COMPUTED TOMOGRAPHY (CT)-SCAN AND METHODS
A corrosion analysis assembly including a computed tomography scanner positioned relative to a component, the computed tomography scanner configured to non-destructively scan a section of the component to identify corrosion sites and measure spatially resolved characteristic parameters for the corrosion sites to provide a corrosion data set.
G01N 23/046 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material using tomography, e.g. computed tomography [CT]
G01B 11/06 - Measuring arrangements characterised by the use of optical techniques for measuring length, width, or thickness for measuring thickness
G01N 17/00 - Investigating resistance of materials to the weather, to corrosion or to light
G01N 21/17 - Systems in which incident light is modified in accordance with the properties of the material investigated
G01N 21/88 - Investigating the presence of flaws, defects or contamination
A defect depth estimation system includes a training system and an imaging system that performs defect depth estimation from a monocular 2D image without using a depth sensor. The training system repeatedly receives a first type of image having a defect, and a second type of image that captures the target object having the defect and provides ground truth data indicating an actual depth of the defect. The training system transforms the first domain and the second domain into a target third domain that reduces a domain gap and trains a machine learning model to learn the actual depth of the defect using the target third domain. The imaging system receives a 2D test image in the first forma and uses the trained machine learning model to determine an estimation of the actual depth of the actual defect and to output estimated the estimation of the actual depth.
A main injector can include multiple sub-element mixers. A first sub-element mixer includes a first main air nozzle circumscribing a first main fuel nozzle. A second sub-element mixer includes a second main air nozzle circumscribing a second main fuel nozzle. An annular combustor can include a circumferential array of main injectors disposed proximate a pilot injector. Each main injector of the array of main injectors can be oriented with first sub-element mixers proximate to the pilot injector and second sub-element mixers spaced axially downstream relative to first sub-element mixers in a staged configuration.
A compressor casing is provided. The compressor casing includes an outer wall, a rail extending inwardly from the outer wall and comprising scallop features encompassing pathways and an inner wall connected with an inboard end of the rail. The inner wall includes a first platform surface at a first side of the rail and including first fillets interfacing with first sides of the scallop features and a second platform surface outboard of the first platform surface at a second side of the rail opposite the first side and including second fillets interfacing with second sides of the scallop features.
A turbine vane for use in a gas turbine engine includes an airfoil section having a concave sidewall and a convex sidewall. Both the concave sidewall and convex sidewall extend spanwise between a platform and a radially outward airfoil tip and chordwise between a leading edge and a trailing edge. The concave sidewall includes a convex ablative region.
Embodiments of the present disclosure generally relate to aircraft engines and, more particularly, to detecting defects in aircraft engines using visual neuromorphic sensors. In some embodiments, an event associated with a portion of an aircraft engine may be identified based on a change on a visual data characteristic from a visual neuromorphic sensor. In response to identifying the event associated with the portion of the aircraft engine, synchronous data from a synchronous data collection sensor coupled to the aircraft engine may be retrieved for a predetermined period of time, and a defect associated with the aircraft engine detected based on the identified event and the synchronous data. Other embodiments may be disclosed or claimed.
A gas turbine engine component having a substrate; a thermal barrier coating on the substrate having a porous microstructure; and a reflective layer conforming to the porous microstructure of the thermal barrier coating, wherein the reflective layer comprises a conforming nanolaminate defined by alternating layers of platinum group metal materials selected from the group consisting of platinum group metal-based alloys, platinum group metal intermetallic compounds, mixtures of platinum group metal with metal oxides and combinations thereof. A capping layer can be added over the reflective layer. A supporting layer can be added between the reflective layer and the thermal barrier coating. A process is also disclosed.
Examples described herein provide a computer-implemented method that includes providing the hybrid electric engine, the hybrid electric engine having a gas generating core and an electric machine powered by electric energy. The method further includes determining, by a processing device, whether a use of the electric energy will increase time on wing of the hybrid electric engine of the aircraft a threshold amount. The method further includes, responsive to determining that the use of energy will increase time on wing the threshold amount, apportioning the electric energy from a battery system of the aircraft to increase the time on wing.
B64D 31/06 - Initiating means actuated automatically
B60L 50/60 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries
B64D 27/02 - Aircraft characterised by the type or position of power plants
B64D 27/10 - Aircraft characterised by the type or position of power plants of gas-turbine type
B64D 27/24 - Aircraft characterised by the type or position of power plants using steam or spring force
A method of fabricating a fibrous ceramic preform includes forming a plurality of coated ceramic tows by introducing a polymeric composition to each of a plurality of ceramic tows, and dispersing and stabilizing each of the plurality of ceramic tows such that a coating of a first polymer binder forms on individual filaments of respective ones of the plurality of ceramic tows, incorporating the plurality of coated ceramic tows into a ceramic fabric, applying a tackifier composition comprising a second polymer binder and a second solvent to the ceramic fabric, and incorporating the ceramic fabric into the preform. The first polymer binder is insoluble in the second solvent.
An aircraft includes a gas turbine engine and an optically-based contrail control system. The gas turbine engine is configured to ingest a first mass flow and to exhaust a second mass flow. The optically-based contrail control system is configured to determine an amount of scattered energy contained in the second mass flow and to determine flow characteristics of the second mass flow based at least in part on molecular components contained in the second mass flow. The optically-based contrail control system determines a level of emissions exhausted from the gas turbine engine based at least in part on a combination of the amount of scattered energy and the flow characteristics.
A turbine component includes a body having a pair of spaced walls, with at least one of the walls for facing a fluid flow when mounted in a gas turbine engine. There are a plurality of wall cooling passages having a generally boomerang shape such that a peak apex is spaced from the wall and an indent apex is adjacent to the wall, with the plurality of wall cooling passages having interior sides extending from the peak apex toward the wall to define a corner. Outer sides extend from the corners with a component away from the wall and to the indent apex. A gas turbine engine is also disclosed.
A shaft bearing retainer assembly is provided that includes an axially extending shaft, a bearing, and a bearing retainer subassembly. The shaft has first and second radial surfaces, a distal end, a bearing seat, and a retainer cavity. The bearing seat is engaged with the first radial surface and extends axially inward from the distal end. The retainer cavity is disposed in the second radial surface of the shaft and extends axially inward from the distal end. The shaft includes a first threaded surface portion disposed in the retainer cavity and a second threaded surface portion in the second radial surface. The bearing has a race mounted in the bearing seat. The bearing retainer subassembly includes first and second retainer rings. The first retainer ring is in threaded engagement with the first threaded surface portion. The second retainer ring is in threaded engagement with the second threaded surface portion.
An inspection system for a gas turbine engine includes a team of terrestrial drones each equipped with at least one inspection sensor and at least one processor. The processor is configured to choreograph operation of the terrestrial drones to each move along an associated drone-specific inspection path and collectively traverse an area of interest in a gas turbine engine; and operate the inspection sensor of each of the terrestrial drones to collect inspection data along the associated drone-specific inspection path.
G05D 1/02 - Control of position or course in two dimensions
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G05D 1/00 - Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
21.
GEARED ARCHITECTURE GAS TURBINE ENGINE WITH PLANETARY GEAR OIL SCAVENGE
A fan drive gear system for a turbofan engine according to an exemplary embodiment of this disclosure, among other possible things includes a sun gear that is rotatable about an axis, a plurality of intermediate gears driven by the sun gear, and a baffle that is disposed between at least two of the plurality of intermediate gears for defining a lubricant flow path from an interface between the sun gear and at least one of the plurality of intermediate gears. The baffle includes a channel with at least one ramp portion directing lubricant.
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
F16H 1/28 - Toothed gearings for conveying rotary motion with gears having orbital motion
F16H 57/08 - General details of gearing of gearings with members having orbital motion
22.
GAS TURBINE ENGINE WITH ENTRAINED PARTICLE SEPARATION SYSTEM
A turbine engine is provided that includes a fan section, a nose cone, a compressor inlet, a fan duct, and a particle separation system. The fan section has a plurality of fan blades disposed around a circumference of the fan section. The nose cone is disposed forward of the fan section. The compressor inlet is disposed aft of the fan section. The fan duct is disposed radially outside of the compressor inlet. The particle separation system is configured to produce an electrostatic charge on one or more surfaces forward of or adjacent to the compressor inlet. The electrostatic charge is configured to cause charged particles present within an air flow entering the engine to divert from the compressor inlet and enter the fan duct.
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
23.
PARTICULATE SEPARATOR ASSEMBLY FOR A GAS TURBINE ENGINE
An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.
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 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
24.
GAS TURBINE ENGINE WITH ENTRAINED PARTICLE AGGLOMERATORS AND METHOD
A turbine engine having an axial centerline is provided that includes a compressor section, a combustor section, an outer casing, an inner diffuser case, a turbine section, a particle separator, and a particle agglomerator. The outer casing is disposed radially outside of and spaced apart from an annular combustor. A diffuser outer diameter (OD) flow path is disposed radially between the outer casing and the outer combustor wall. The inner diffuser case is disposed radially inside of and spaced apart from the annular combustor. A diffuser inner diameter (ID) flow path is disposed radially between the inner combustor wall and the inner diffuser case. The particle agglomerator is configured to produce acoustic signals that causes agglomeration of particles entrained in an air flow within the turbine engine.
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
F01D 25/32 - Collecting of condensation waterDrainage
25.
MULTI-PHASE RADIATIVE AND THERMAL BARRIER COATING SYSTEM
A multi-phase radiative and thermal barrier coating system including a substrate having a substrate surface; a first layer deposited on the substrate surface; a second layer deposited on the first layer, the second layer having an outer surface, wherein the second layer comprises radiative barrier materials in a porous thermal conduction and radiant heat transfer resistant microstructure.
An airfoil includes an airfoil section that defines a trailing edge region that includes a trailing edge. The airfoil section is formed of a ceramic matrix composite that includes core fiber plies and skin fiber plies. The core fiber plies define a radial tube that has a radiused end in the trailing edge region. The skin fiber plies wrap around the core fiber plies and a filler element in the trailing edge region is aft of the internal cavity and is sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. There is at least one cooling passage that includes a first, inlet orifice section that opens to the internal cavity at a location forward of the radiused end and that extends through the core fiber plies, and a second, outlet orifice section that extends through the trailing edge.
A CMC airfoil includes core fiber plies that define a radial tube that circumscribes an internal cavity, and skin fiber plies that define an exterior of the airfoil section. There is a filler element in the trailing edge region aft of the internal cavity and sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. At least one cooling passage includes at least one inlet orifice section that opens to the internal cavity and extends through the core fiber plies, a radially-elongated multi-dimensional plenum connected with the at least one inlet orifice section, a plurality of outlet orifice sections extending through the trailing edge, and a plurality of intermediate passage sections bound by at least one of the skin plies and the filler element and connecting the radially-elongated multi-dimensional plenum with the plurality of outlet orifice sections.
A gas turbine engine combustor panel assembly includes a shell defining an outer periphery of a combustion chamber with an outer skin with a first plurality of holes, a combustor panel including a first end, a second plurality of holes extending through an inner face to an inner void, to an outer face with a third plurality of holes to allow fluid flow from the first plurality of holes to the combustion chamber, a plurality of bolt holes, a second end with an internal combustion chamber section extending axially from a central axis and from a panel wall of the combustor panel. The assembly includes a sheet metal seal plate abutting the plurality of bolt holes and the interface, and a plurality of bolts extending along a central axis and interfacing with the plurality of bolt holes to connect the combustor panel and shell.
A machine has: an outer member; an inner member having an outer diameter (OD) surface; and a seal system. A seal housing is mounted to the outer member and has first and second walls; a first seal stage contacting the OD surface and the first wall; a second seal stage contacting the OD surface and the second wall; and a wave spring biasing the seal stages axially apart from each other. The seal stages each have a plurality of seal segments interfitting end to end and each having: a first end; a second end circumferentially opposite the first end; a first face; a second face axially opposite the first face; an inner diameter (ID) face; and an outer diameter (OD) face having a radially protruding lug. The housing has an inner diameter (ID) surface having recesses receiving the lugs of the seal stages to circumferentially retain the seal stages.
A power system includes a power source configured to output electrical power, a power converter configured to convert the electrical power into a converted power, and a power bus configured to deliver the converted power to a power load connected to the power bus. The power system further includes a controller that implements a neural network (NN) trained to perform a NN-based model predictive control (NNMPC). The controller utilizes the NNMPC to obtain at least one learned control input for power regulation for a current system state and power load measurement of the power load in real-time, and to perform an output action that regulates the power system based on the at least one learned control input obtained by the NNMPC.
B60L 58/10 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries
G05B 13/02 - Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
G05B 13/04 - Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
H02J 1/08 - Three-wire systemsSystems having more than three wires
31.
SELECTIVE POWER DISTRIBUTION FOR AN AIRCRAFT PROPULSION SYSTEM
An aircraft assembly includes a geartrain, a first propulsor rotor and a rotating assembly. The geartrain includes a first gear system and a second gear system. The first gear system includes a first sun gear, a first ring gear, a plurality of first intermediate gears and a first carrier. The first intermediate gears are radially between and meshed with the first sun gear and the first ring gear. The second gear system includes a second sun gear, a second ring gear, a plurality of second intermediate gears and a second carrier. The second sun gear is rotatable about the axis with the first sun gear. The second intermediate gears are radially between and meshed with the second sun gear and the second ring gear. The second carrier is coupled to the first ring gear. The rotating assembly is configured to drive rotation of the first propulsor rotor through the geartrain.
An engine assembly is provided that includes a sun gear, a ring gear, a plurality of intermediate gears, a carrier, a first rotating structure, a second rotating structure and a first bearing. The ring gear is rotatable about an axis and circumscribes the sun gear. The intermediate gears are arranged circumferentially about the axis in an array. Each of the intermediate gears is radially between and meshed with the sun gear and the ring gear. The carrier is rotatable about the axis. Each of the intermediate gears is rotatably mounted to the carrier. The first rotating structure is configured as or otherwise includes the carrier. The second rotating structure is configured as or otherwise includes the ring gear. The first bearing is radially between and engaged with the first rotating structure and the second rotating structure.
An electronically driven lubrication system including an electric lubrication pump; at least one bearing for a gas turbine engine component; an oil cooler fluidly coupled with the electric lubrication pump; lubrication oil fluidly coupled with the electric lubrication pump, at least one bearing and oil cooler; at least one sensor in operative communication with the lubrication oil; a controller comprising a processor in operative communication with the at least one sensor, the electric lubrication pump, the at least one bearing and the oil cooler; and the processor configured to provide processor outputs to the electric lubrication pump responsive to data collected from the at least one sensor, wherein the processor employs an operational configuration for the electronically driven lubrication system.
F02C 7/14 - Cooling of plants of fluids in the plant
F16N 29/00 - Special means in lubricating arrangements or systems providing for the indication or detection of undesired conditionsUse of devices responsive to conditions in lubricating arrangements or systems
34.
SYSTEMS AND METHODS FOR IMPROVING BACKWARD FLOW FORMING OF SHAFTS
An apparatus for backward flow forming a material may comprise a mandrel having a headstock at a proximate end of the mandrel, the mandrel configured to rotate about an axis, a plurality of rollers disposed radially outward of the mandrel configured to exert force on the material to form a work piece at a plastic deformation zone, wherein the work piece flows from the plastic deformation zone between the plurality of rollers and the mandrel toward a distal end of the mandrel, and a catcher, coaxial to the mandrel, and removably coupled to the work piece at a traveling end of the work piece.
An assembly is provided for an aircraft. This aircraft assembly includes a first propulsor rotor, a geartrain, a rotating assembly and an auxiliary turbine. The rotating assembly is rotatable about an axis and includes a turbine rotor. The rotating assembly is coupled to and is configured to drive rotation of the first propulsor rotor through the geartrain. The auxiliary turbine is coupled to the first propulsor rotor independent of the geartrain.
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
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 7/32 - Arrangement, mounting, or driving, of auxiliaries
36.
FLUID DAMPER FOR TURBINE ENGINE GEARTRAIN ASSEMBLY
An engine assembly is provided that includes a geartrain, a device and a fluid damper. The geartrain is configured as or otherwise includes an epicyclic gear system. A first component of the geartrain is rotatable about an axis. The device is configured to brake and/or lock rotation of the first component of the geartrain about the axis. The device is configured as or otherwise includes a device structure. The fluid damper engages the device structure. The fluid damper is configured to damp vibrations in the device.
A sensor arrangement for a gas powered turbine includes at least one sensor disposed on a power extraction shaft and configured to output a measured power extraction of the power extraction shaft. A controller is in communication with the at least one sensor. The controller includes a memory and a processor. The memory stores instructions for causing the processor to respond to a received measured power extraction of the power extraction shaft by synthesizing an instantaneous engine power output and engine efficiency and adjusting at least one parameter of the engine based on the synthesized engine power output and engine efficiency.
A method applies one or more films of polynuclear aluminum oxide hydroxide and polynuclear chromium hydroxide to a metal substrate. A method thermally treats the metal substrate with the one or more films at a temperature of at least 250° C., the thermal treatment reducing the polynuclear aluminum oxide hydroxides and the polynuclear chromium hydroxides to at least one layer of aluminum-chromium oxide.
F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coatingContact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
C23C 24/08 - Coating starting from inorganic powder by application of heat or pressure and heat
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
C23C 30/00 - Coating with metallic material characterised only by the composition of the metallic material, i.e. not characterised by the coating process
C25D 13/18 - Electrophoretic coating characterised by the process using modulated, pulsed or reversing current
An assembly is provided for an aircraft powerplant. This assembly includes a differential geartrain, a first rotating assembly, a second rotating assembly, a first actuator and a second actuator. The differential geartrain includes a sun gear, a ring gear, a plurality of intermediate gears and a carrier. The first rotating assembly is coupled to the differential geartrain through the carrier. The first rotating assembly includes a first turbine rotor. The second rotating assembly is coupled to the differential geartrain through the ring gear. The second rotating assembly includes a second turbine rotor. The first actuator is coupled to the differential geartrain through the ring gear. The second actuator is coupled to the differential geartrain through the sun gear.
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
F16H 3/72 - Toothed gearings for conveying rotary motion with variable gear ratio or for reversing rotary motion using gears having orbital motion with a secondary drive, e.g. regulating motor, in order to vary speed continuously
F16H 48/10 - Differential gearings with gears having orbital motion with orbital spur gears
40.
HYDROGEN STEAM INJECTED TURBINE ENGINE WITH TURBOEXPANDER HEAT RECOVERY
A propulsion system for an aircraft includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with a gaseous fuel and ignited to generate an exhaust gas flow that is expanded through a turbine section. A fuel system supplies a fuel to the combustor through a fuel flow path, a first heat exchanger thermally communicates a first heat load into a cooling flow, a turboexpander where a heated cooling flow from the first heat exchanger is expanded to generate shaft power and cooled to provide a cooled cooling flow, and a second heat exchanger thermally communicates a second heat load to the cooled cooling flow that is communicated from the turboexpander, cooling flow from the second heat exchanger is communicated to the combustor section.
F02C 7/141 - Cooling of plants of fluids in the plant of working fluid
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
41.
AUXETIC MATERIALS AND STRUCTURES FOR CERAMIC MATRIX COMPOSITE AIRFOIL MANDRELS
A method of fabricating a ceramic matrix composite component includes fabricating a ceramic preform, the preform comprising a hollow portion with an internal cavity extending along a first axis from a first end to a second end of the hollow portion, supporting the hollow portion with an auxetic mandrel disposed within and the internal cavity and coaxial with the hollow portion, the auxetic mandrel comprising a first mandrel end and a second mandrel end, at least partially densifying the preform with a matrix, and removing the auxetic mandrel from the internal cavity by simultaneously applying a compressive force to each of the first mandrel end and the second mandrel end creating a deformed auxetic mandrel to reduce a cross-sectional profile of the auxetic mandrel along a second axis orthogonal to the first axis, and extracting the deformed auxetic mandrel from the first end of the hollow portion.
An assembly is provided for a turbine engine. This assembly includes a compressor rotor and a flowpath wall. The compressor rotor is rotatable about an axis. The compressor rotor includes a plurality of compressor blades arranged circumferentially around the axis. The flowpath wall forms an outer peripheral boundary of a flowpath in which the compressor blades are disposed. The flowpath wall includes a polymer shell and a metal liner bonded to the polymer shell. The polymer shell axially overlaps and circumscribes the metal liner. The metal liner axially overlaps and circumscribes the compressor blades.
An assembly for a combustor for a gas turbine engine, including: a plurality of heat shield panels attached to at least one combustor liner, each one of the plurality of heat shield panels including a panel portion and a first forward rail and a second rearward rail each extending from an outer surface of the panel portion, the panel portion including an inner surface opposite of the outer surface, the panel portion also includes a forward end and a rearward end, the forward end of the panel portion is axially forward of the rearward end of an adjacent heat shield panel of the plurality of heat shield panels such that a gap is defined between the outer surface of the panel portion of one heat shield panel of the plurality of heat shield panels and an inner surface of the panel portion of an adjacent heat shield panel of the plurality of heat shield panels, the panel portion also includes a plurality of apertures extending from the outer surface to the inner surface; and a plurality of cooling pins extending upwardly and away from the outer surface towards a surface of the at least one combustor liner.
A turbine section for a gas turbine engine includes a turbine having at least one blade rotatable around an axis. The at least one blade has a tip. The turbine section for a gas turbine engine also includes at least one blade outer air seal arranged radially outward from the tip. The blade outer air seal has a center web and first and second mounting arms extending from the center web. Each of the first and second mounting arms include at least one aperture configured to receive a pin to attach the blade outer air seal to an engine static structure. A gas turbine engine and a method of attaching a blade outer air seal to a static structure of a gas turbine engine are also disclosed.
A stator assembly is provided and includes a stator element and a radial height adjustment mechanism. The stator assembly includes an inboard portion which establishes a primary clearance with rotor elements and exhibits a measurable parameter corresponding to the primary clearance and an outboard portion integrally formed with the inboard portion. The radial height adjustment mechanism is coupled with the outboard portion and configured to be operable, based on the measurable parameter, to adjust a radial height of the stator element and in turn to adjust the primary clearance.
A user interface for a lightweight viewer may include, among other things, a viewing window operable to display geometry established by one or more geometric objects of a tessellated model. The one or more geometric objects include one or more visual attributes that depict the respective geometry. A palette menu is established by a plurality of menu objects of the tessellated model. The menu objects are assigned respective visual attribute settings. The menu objects are associated with the one or more geometric objects such that selection of the respective menu object causes at least one of the one or more visual attributes of the associated one or more geometric objects to update in the viewing window according to the respective visual attribute setting. A method of establishing a tessellated model is also disclosed.
G06F 3/04845 - Interaction techniques based on graphical user interfaces [GUI] for the control of specific functions or operations, e.g. selecting or manipulating an object, an image or a displayed text element, setting a parameter value or selecting a range for image manipulation, e.g. dragging, rotation, expansion or change of colour
G06F 3/0482 - Interaction with lists of selectable items, e.g. menus
G06T 17/20 - Wire-frame description, e.g. polygonalisation or tessellation
A user interface for a lightweight viewer may include, among other things, a viewing window operable to display geometry established by one or more geometric objects of a tessellated model. The one or more geometric objects are associated with respective edges of the geometry. One or more annotation objects of the tessellated model are assigned information associated with the respective edges and are operable to depict the respective information as a graphical annotation in the viewing window. One or more curve objects of the tessellated model are associated with the respective one or more annotation objects and include respective curves dimensioned to follow the respective edges of the geometry. The one or more curve objects are operable to selectively depict the respective edges of the geometry in the viewing window in response to selection of the respective annotation object. A method of establishing a tessellated model is also disclosed.
G06F 30/12 - Geometric CAD characterised by design entry means specially adapted for CAD, e.g. graphical user interfaces [GUI] specially adapted for CAD
G06T 17/20 - Wire-frame description, e.g. polygonalisation or tessellation
48.
ANTI-ROTATION FEATURE FOR INSTRUMENTATION OF GAS TURBINE ENGINE
A component assembly of a gas turbine engine includes a component of a gas turbine engine, and an instrumentation probe installed to the component. The instrumentation probe includes a sensor body extending through a component wall, and a threaded fastener installed onto a complimentary thread of the sensor body to retain the sensor body. The threaded fastener is retained to the sensor body via deforming a threaded interface between the threaded fastener and the sensor body. A method of installing an instrumentation probe to a component of a gas turbine engine includes installing a sensor assembly of the instrumentation probe through a component wall, securing a threaded fastener to a complimentary thread of the sensor assembly to retain the sensor body, and deforming a threaded interface between the threaded fastener and the sensor assembly to prevent rotation of the threaded fastener relative to the sensor assembly.
An aluminum containing component comprises an aluminum alloy and an aluminum oxide layer disposed on the aluminum alloy. The aluminum oxide layer comprises crystalline aluminum oxide. The aluminum containing component is at least one of vane, a fan blade or a fan casing of a low pressure compressor section of a gas turbine. In an embodiment, a method comprises disposing an aluminum containing component in an electrochemical cell that comprises a dilute alkaline solution. The aluminum containing component is electrically contacted to become a first electrode in the electrochemical cell. The wall of the bath is electrically contacted to act as a second electrode in the electrochemical cell. A voltage is applied between the first electrode and the second electrode to form an aluminum oxide layer on the aluminum containing component.
A method for coating a component having: a metallic substrate; a ceramic coating having one or more ceramic coating layers atop the substrate; and a cooling passageway system comprising a plurality of feed passageways extending from one or more inlet ports and a plurality of outlet passageways. The outlet passageways have openings in the coating. The method involves: applying a slurry aluminide to the plurality of outlet passageways; coupling the one or more inlets to a suction source; applying an external gas flow to the component, the suction source drawing the external gas in through the outlet passageways and out through the one or more inlet ports, the external gas flow comprising at least 50% by volume combined one to all of Ar, He, and H2; and while the suction source is drawing the external gas, heating the component to aluminize the cooling passageway system.
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
C23C 24/10 - Coating starting from inorganic powder by application of heat or pressure and heat with intermediate formation of a liquid phase in the layer
F01D 5/18 - Hollow bladesHeating, heat-insulating, or cooling means on blades
An apparatus is provided for a turbine engine. This turbine engine apparatus includes an annular nozzle. The annular nozzle includes an inner monolithic body and an outer monolithic body. The inner monolithic body includes an inner shroud and a plurality of vanes. The inner shroud extends axially along and circumferentially around an axis. The vanes are arranged circumferentially about the axis in an array. Each of the vanes projects radially out from the inner shroud to a respective outer distal end. The outer monolithic body is radially outboard of and circumscribes the inner monolithic body. The outer monolithic body is configured as or otherwise includes an outer shroud. The outer shroud extends axially along and circumferentially around the axis. The outer shroud radially engages each of the vanes at the respective outer distal end.
A core section and nacelle assembly of a gas turbine engine includes a compressor located at an engine central longitudinal axis, a core case enclosing the compressor, and a nacelle located radially outboard of the core case and defining a core compartment between the nacelle and the core case. One or more vent openings are located in the nacelle to circulate a cooling airflow through the core compartment, and one or more fans are positioned at the one or more vent openings to urge the cooling airflow through the one or more vent openings to cool the core compartment.
An assembly is provided for a turbine engine. This turbine engine assembly includes a supplemental thrust section and a duct. The supplemental thrust section includes a rotating detonation combustor. The duct includes a supplemental thrust section inlet fluidly coupled with and leading to the rotating detonation combustor. The supplemental thrust section inlet has a flow area that decreases as at least a first portion of the supplemental thrust section inlet extends towards the rotating detonation combustor.
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
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
A fan drive gear system for a turbine engine includes a sun gear that is configured to be driven by an engine shaft that is rotatable about an axis, a plurality of intermediate gears that are intermeshed with the sun gear, a ring gear assembly that is engaged with the plurality of intermediate gears, the ring gear is configured for attachment to a static structure, a carrier that supports rotation of the plurality of intermediate gears, the carrier configured for rotation about the axis, at least one baffle that is attached to the carrier that is configured to impart a momentum on expelled lubricant, a fixed gutter that is disposed radially outside the at least one baffle and is configured to receive lubricant that is exhausted from the at least one baffle, and a fan shaft that is rotatable about the axis and configured to be driven by the carrier.
A user interface for a lightweight viewer may include, among other things, a viewing window operable to display geometry established by one or more geometric objects of a tessellated model and an information window established by a plurality of content objects of the tessellated model. The content objects are associated with respective layers of the tessellated model that occupy a common display region. The content objects are operable to selectively display information associated with the tessellated model in response to user interaction with the respective content object such that the respective layer is activated but a remainder of the layers are deactivated in the common display region. A method of establishing a tessellated model is also disclosed.
G06F 30/12 - Geometric CAD characterised by design entry means specially adapted for CAD, e.g. graphical user interfaces [GUI] specially adapted for CAD
G06F 3/0483 - Interaction with page-structured environments, e.g. book metaphor
A user interface for a lightweight viewer may include, among other things, a viewing window operable to display geometry established by one or more of geometric objects of a tessellated model. One or more annotation objects of the tessellated model are assigned information associated with the respective one or more geometric objects and are operable to depict the respective information as a graphical annotation in the viewing window. A search window is established by a plurality of content objects of the tessellated model. The content objects include a query object and a results object dynamically linked to the query object. The query object establishes a query input field operable to query the information of the one or more annotation objects. The results object is operable to display a list of results meeting the query. A method of establishing a tessellated model is also disclosed.
G06F 30/12 - Geometric CAD characterised by design entry means specially adapted for CAD, e.g. graphical user interfaces [GUI] specially adapted for CAD
G06F 3/0482 - Interaction with lists of selectable items, e.g. menus
G06T 11/60 - Editing figures and textCombining figures or text
G06T 17/20 - Wire-frame description, e.g. polygonalisation or tessellation
A user interface for a lightweight viewer may include, among other things, a viewing window operable to display geometry established by one or more of geometric objects of a tessellated model and a plurality of notification objects of the tessellated model. The notification objects include a first statement object and an acknowledgement object linked to the first statement object. The first statement object is operable to establish a first notification screen that blocks display of the geometry in the viewing window prior to selection of the acknowledgement object but permits display of the geometry in response to selection of the acknowledgement object. A method of establishing a tessellated model is also disclosed.
An embedded processing system includes processing circuitry, a memory system, and a plurality of attached modular components. The attached modular components are each provided with a nameplate including at least part and serial number data. The processing circuitry is operable to receive the nameplate information from each of the attached modular components and compare the received nameplate information with stored nameplate information for the particular attached modular component. The processing circuitry is operable to communicate with the attached modular component if the received nameplate information matches the stored nameplate information and identify a fault if the received nameplate information conflicts with the stored nameplate information. A method and an assembly are also disclosed.
G05B 19/4155 - Numerical control [NC], i.e. automatically operating machines, in particular machine tools, e.g. in a manufacturing environment, so as to execute positioning, movement or co-ordinated operations by means of programme data in numerical form characterised by programme execution, i.e. part programme or machine function execution, e.g. selection of a programme
An embedded processing system and access combination includes processing circuitry, a memory system, and a plurality of user credential files. The user credential files include an encrypted user identifier, and an encrypted list of authorized task roles the particular user would have within the embedded processing system. Expected credentials from the user credential files are stored in the memory system. The processing system is programmed to receive a user credential file from a user, and compare expected credentials within the memory system to identify if the user is an authorized user. The processing system is programmed to allow access to an authorized user and deny access to an unauthorized user and determine what task roles are authorized for the authorized user, and deny access for the authorized user to other tasks. A method and an assembly are also disclosed.
A turbine engine airfoil element has a plurality of main body passageways along a camber line and having two or more camberwise/chordwise distributed protuberant portions with necks between the protuberant portions. A plurality of skin passageways include: at least one first skin passageway each nested between a first of the pressure side and the suction side and two adjacent main body passageways; and a plurality of second skin passageways each nested between first of the pressure side and the suction side and two said protuberant portions of a corresponding main body passageway.
A fan drive gear system for a turbine engine includes an auxiliary reservoir that is disposed radially outward of a rotating carrier and an oil director that is attached to the carrier. The oil director imparts a rotational flow direction into expelled oil to drive the expelled oil tangentially against an oil receiving surface and into the auxiliary reservoir.
A stator cluster is provided and includes inner and outer stator walls, stator vanes radially interposed between the inner and outer stator walls and a stanchion body connected to and extending radially outwardly from the outer stator wall. At least the stator vanes, the outer stator wall and the stanchion body are formed to define internal paths.
A rear bearing air cooling system including an engine inlet section housing a nose cone; a central rotor shaft interior of the nose cone; a rear bearing supporting the central rotor shaft; and a cooling air passage formed through the nose cone and the central rotor shaft, the cooling air passage in fluid communication with the nose cone, the rear bearing and a nose cone vent.
Vane assemblies and gas turbine engines having vane assemblies are described. The vane assemblies include a vane having outer and inner diameter ends and at least a leading-edge cavity therein. A direction of flow through the leading-edge cavity is from the outer diameter end toward the inner diameter end. An inner diameter platform is arranged at the inner diameter end of the vane and includes an inner diameter flow path having an exit at an aft side of the inner diameter platform. The inner diameter platform includes an inner diameter flow aperture fluidly coupling the leading-edge cavity and the inner diameter flow path. A baffle is installed within the leading-edge cavity and arranged to reduce a cross-sectional area in the direction of flow of the leading-edge cavity such that the leading-edge cavity has a larger cross-sectional area proximate the outer diameter end than at the inner diameter end.
A stator assembly of a gas turbine engine includes a stator including at least a stator vane and a stator inner platform located radially inboard of the stator vane. A stator inner airseal is positioned radially inboard of the stator inner platform. The stator inner airseal is configured to define a seal arrangement with two or more knife edges of a radially adjacent rotating component. A sensor is positioned at the stator inner airseal. The sensor includes a sensor body configured to be positioned axially between the knife edges, and configured to detect a radial distance from the sensor to the radially adjacent rotating component.
A thermal spray powder includes composite powder particles that each have a silicon carbide core and a multi-layered shell that surrounds the core. The shell includes at least one layer of silica and at least one layer of alumina. The powder is used to deposit a mullite-based topcoat on a ceramic matrix composite wall of an article.
C04B 35/622 - Forming processesProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products
C04B 35/565 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxides based on carbides based on silicon carbide
A method for detecting performance degradation of an AC Nozzle in a gas turbine engine including operating the gas turbine engine under a normal operating condition, maintaining the fuel flow rate at a constant fuel flow rate, determining a delta pressure PDelta-Normal across an N number of Proportional Metering Valves (PMVs) as the gas turbine engine is operating under the normal operating condition, selecting a PMV and an AC Nozzle, controlling the PMV to modify a fuel flow rate to the selected AC Nozzle to cause the gas turbine engine to operate under a modified fuel flow condition, determining a delta pressure PDelta-Modified across the N number of PMVs as the gas turbine engine is operating under the modified operating condition and comparing the delta pressure PDelta-Normal with the delta pressure PDelta-Modified to determine if the selected AC Nozzle is uncalibrated or has failed.
A grid arrayed microtube heat exchanger with vibration dampening support including an upper portion comprising an upper portion support wall having multiple upper portion receivers; a lower portion comprising a lower portion support wall having multiple lower portion receivers; a grid array comprising multiple rows of the lower portion receivers and the upper portion receivers; multiple microtubes supported by the upper portion receivers and the lower portion receivers; a gap located between each microtube; and a support insertable through the gap between the multiple microtubes, the support including at least one cam contacting the microtube, the at least one cam being rigid.
F28D 7/16 - Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall the conduits being arranged in parallel spaced relation
A component having a thermal barrier coating having a substrate surface of the component; a thermal barrier coating on the substrate surface, the thermal barrier coating having a coating surface and coating surface accessible spaces; and a layer of rare-earth phosphate on surfaces of the coating surface accessible spaces. The spaces can be intercolumnar spaces between columns or feathers of the thermal barrier coating. A method is also disclosed.
C23C 16/02 - Pretreatment of the material to be coated
C23C 16/455 - Chemical coating by decomposition of gaseous compounds, without leaving reaction products of surface material in the coating, i.e. chemical vapour deposition [CVD] processes characterised by the method of coating characterised by the method used for introducing gases into the reaction chamber or for modifying gas flows in the reaction chamber
70.
GAS TURBINE ENGINE BEARING AND GEARBOX ARRANGEMENT
A gas turbine engine includes a plurality of blades that deliver air into a compressor section and a drive gear system including an epicyclic transmission. A drive shaft interconnects a gear carrier of the epicyclic transmission and the plurality of blades. A turbine section drives an input of the drive gear system through a turbine shaft. The drive gear system is straddle mounted by a first bearing forward of the drive gear system and a second bearing aft of the drive gear system. A ring gear of the epicyclic transmission is coupled to an engine case with a compliant flexure at a position between the first bearing and the second bearing.
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
71.
CONDENSER FOR HYDROGEN STEAM INJECTED TURBINE ENGINE
A propulsion system for an aircraft includes a hydrogen fuel system supplying hydrogen fuel to the combustor through a fuel flow path. A condenser extracts water from an exhaust gas flow and includes a plurality of spiral passages disposed within a collector. The spiraling passages generate a transverse pressure gradient to direct water out of the exhaust gas flow toward the collector.
A turbine engine assembly generates an exhaust gas flow that is divided into a first exhaust gas flow and a second exhaust gas flow. A desiccation system transfers water vapor from the first exhaust gas flow into the second exhaust gas flow. A condenser extracts water from the second exhaust gas flow and an evaporator system uses thermal energy from the exhaust gas flow to generate a steam flow from at least a portion of water that is extracted by the condenser for injection into the core flow path.
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F02C 3/34 - Gas-turbine plants characterised by the use of combustion products as the working fluid with recycling of part of the working fluid, i.e. semi-closed cycles with combustion products in the closed part of the cycle
A method for protecting a coated substrate having a porous ceramic barrier coating includes applying a molten salt to the ceramic barrier coating. The salt is selected from the group consisting of one or more acetates and/or nitrates. The molten salt is infiltrated into porosity of the ceramic barrier coating. The infiltrated molten salt is solidified. The solidified salt is sintered.
A disclosed turbine engine having a fan drive gear system includes a first lubrication circuit having a first pump communicating oil to at least the fan drive gear system and at least one low shaft bearing assembly. The first pump is driven by the low spool through a first tower shaft. A second lubrication circuit includes a second pump driven by one of a high spool and the low spool separate from the first pump and communicates oil to engine components separate from the first lubrication system.
An aircraft propulsion system includes a condenser at least partially disposed within the core flow path where water is extracted from the exhaust gas flow, an evaporator system that is at least partially disposed within the core flow path upstream of the condenser where thermal energy from the exhaust gas flow is utilized to generate a steam flow from at least a portion of water that is extracted by the condenser. Steam within the exhaust gas flow is concentrated into a portion of the exhaust gas flow that is communicated through the condenser.
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
F01K 3/26 - Plants characterised by the use of steam or heat accumulators, or intermediate steam heaters, therein having heaters with heating by steam
F01K 7/16 - Steam engine plants characterised by the use of specific types of enginePlants or engines characterised by their use of special steam systems, cycles or processesControl means specially adapted for such systems, cycles or processesUse of withdrawn or exhaust steam for feed-water heating the engines being only of turbine type
A method of providing access to flight-related data of aircraft includes correlating a plurality of data sets that are different from each other and each contain flight-related data for one or more aircraft engines, to determine a plurality of flights of the one or more aircraft engines that correspond to the plurality of data sets. At least a portion of the flight-related data describes operation of the aircraft engines during the plurality of flights. The correlating is performed based on metadata of the plurality of data sets, and the plurality of data sets includes a plurality of first data sets that utilize a plurality of different schemas. The method also includes creating, for each of the plurality of flights and based on the correlating, a respective flight object that represents the flight and includes a plurality of discrete flight data objects that each correspond to a respective one of the plurality of data sets for the flight; and utilizing the flight object for one of the flights to provide access to one or more of the data sets for said one of the flights. A system for aircraft data management is also disclosed.
G06F 16/25 - Integrating or interfacing systems involving database management systems
G06F 16/27 - Replication, distribution or synchronisation of data between databases or within a distributed database systemDistributed database system architectures therefor
77.
ARTICLE AND METHOD FOR ACCURATELY POSITIONING OPTICAL TIP-TIMING PROBE
A method for accurately positioning an optical tip-timing probe for identifying vibration in a rotor blade tip of a turbine engine, wherein the method includes identifying a target location to be measured on a rotor blade tip of a plurality of rotor blades disposed within a rotor casing of a turbine engine, marking the target location with a reflective material, associating a probe body having probe head cavity with the rotor casing, such that the probe head cavity is aligned with a rotor casing opening which exposes the plurality of rotor blades, disposing the probe head within the probe head cavity to be movably aligned with the rotor casing opening, centering the probe head within the probe head cavity to be located in a nominal starting position and operating the optical tip-timing probe to identify the target location and precisely align the probe head with the target location.
A turbine engine assembly generates an exhaust gas flow that is communicated through a core flow path. The exhaust gas flow is split into a first exhaust gas flow and a second exhaust gas flow. Water is extracted in a condenser from the second exhaust gas flow. The extracted water is transformed into a steam flow in an evaporator system utilizing thermal energy from at least the second exhaust gas flow. An exit flow from the condenser is communicated through an exhaust compressor and compressed to a higher pressure exit flow.
A turbine engine assembly generates an exhaust gas flow that is communicated through a core flow path. The exhaust gas flow is split into a first exhaust gas flow and a second exhaust gas flow. Water from the second exhaust gas flow is condensed and extracted by a condenser. The extracted water is transformed into a steam flow within an evaporator system utilizing thermal energy from at least the second exhaust gas flow.
F01N 3/18 - Exhaust or silencing apparatus having means for purifying, rendering innocuous, or otherwise treating exhaust for rendering innocuous by thermal or catalytic conversion of noxious components of exhaust characterised by methods of operationControl
F01N 3/24 - Exhaust or silencing apparatus having means for purifying, rendering innocuous, or otherwise treating exhaust for rendering innocuous by thermal or catalytic conversion of noxious components of exhaust characterised by constructional aspects of converting apparatus
F02C 7/143 - Cooling of plants of fluids in the plant of working fluid before or between the compressor stages
A coated knife edge seal member has an annular knife edge having: a flank having a first end face and a second end face. The knife edge has: a tip converging to a rim; and an annular reference datum. The member has a metallic substrate and a coating on the substrate at the tip.
In a method for repairing a coated article, the article has: a ceramic matrix composite (CMC) substrate; and a coating system having a plurality of layers. A damage site at least partially penetrates at least one of the layers. The method includes: applying a slurry of a repair material to the damage site for repairing a first of the penetrated layers; and after the applying, heating the repair material with a plasma torch.
An assembly for an aircraft propulsion system includes a propulsor and an imaging assembly. The propulsor includes a propulsor disk, a plurality of propulsor blades, and a nose cone. The plurality of propulsor blades are circumferentially distributed about the propulsor disk. Each propulsor blade of the plurality of propulsor blades extends radially between and to a root end and a tip end. The root end is disposed at the propulsor disk. The propulsor disk and the plurality of propulsor blades are configured to rotate about the rotational axis. The nose cone is disposed axially adjacent the propulsor disk. The imaging assembly includes an imaging device disposed on the nose cone. The imaging device includes a camera. The camera is configured to capture image data of each propulsor blade of the plurality of propulsor blades as the plurality of propulsor blades rotate about the rotational axis.
A method of improving surface roughness of ceramic matrix composites (CMCs) is provided. The method includes completing a formation of the CMCs and a chemical vapor infiltration (CVI) process to initially coat the CMCs, inspecting a CMC surface, identifying, from a result of the inspecting, a defect in the CMC surface that negatively impacts a surface roughness characteristic thereof, locally targeting and ablating the defect and re-inspecting the CMC surface to ensure that the defect is correct.
An accessible debris separator for a gas turbine engine is provided. The accessible debris separator including a chamber formed by a entrance wall, an exit wall, an inner wall, and an outer wall; a plurality of inlet openings in the entrance wall; a plurality of outlet openings in the exit wall; and a component within the chamber, the component forcing cooling air with debris particulates entering the chamber via an inlet opening of the plurality of inlet openings to take a circuitous path thereby separating the debris particulates from the cooling air prior to the cooling air exiting the chamber via an outlet opening of the plurality of outlet openings.
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
F01D 11/10 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
A vane assembly includes a vane having an outer platform with mount structure and an airfoil extending inwardly from the mount structure and formed of ceramic matrix composites (“CMC”). The vane has an internal cooling chamber. A metal baffle is received within the internal cooling chamber and extends beyond the platform. A bathtub seal has a central portion receiving the metal baffle. The metal baffle has a central chamber and cooling air holes to deliver cooling air into the internal cooling chamber. The bathtub seal has an inner wall and an outer wall outward of the metal baffle. The inner wall, the outer wall and a bottom wall define a bathtub seal chamber. The bottom wall is in contact with an upper surface of the vane. The vane is connected to the outer wall. Static structure is received within the bathtub seal chamber. A gas turbine engine is also disclosed.
A stator vane for a gas turbine engine combustor has a vane segment that includes a vane having a leading edge, a trailing edge, a suction side, a pressure side, a vane inner base, and a vane outer base, an inner shroud segment attached to the vane at the vane inner base, and an outer shroud segment attached to the vane at the vane outer base. The stator vane further includes a support structure attached to the vane at a vane trailing edge, and a first weld shield attached to the support structure by a first connector. The first weld shield is positioned over and spaced away from the suction side. The vane, inner shroud segment, outer shroud segment, support structure, and first weld shield are integrally formed during a single, continuous additive manufacturing process.
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
B22F 7/06 - Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting of composite workpieces or articles from parts, e.g. to form tipped tools
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
A gas turbine engine includes a ceramic matrix composite (CMC) vane arc segment that has first and second platforms and an airfoil section that extends radially therebetween. The airfoil section includes an internal through-cavity, and the first and second platforms include, respectively, platform inlet and outlet ports connected to the internal through-cavity for conveying cooling air. The CMC vane arc segment is radially mounted between first and second metallic vane supports. The second metallic vane support includes a plenum and a plenum inlet port connected with the platform outlet port for receiving the cooling air from the internal through-cavity. There is a seal located radially between the second platform and the second metallic vane support. The seal circumscribes the platform outlet port to limit leakage of the cooling air between the second platform and the second metallic vane support.
An article includes a ceramic matrix composite (CMC) vane arc segment for disposal about an engine central axis. The CMC vane arc segment includes first and second platforms and an airfoil section that extends therebetween. The airfoil section defines a radial stacking axis that is oblique to a Z-axis that is perpendicular to the engine central axis such that the radial stacking axis forms a non-zero angle with the Z-axis.
A method of operation is provided during which heat energy is transferred into a first heat exchange fluid during a first mode of operation and during a second mode of operation. The first heat exchange fluid is directed through a first flowpath of a heat exchanger during the first mode of operation and during the second mode of operation. A second heat exchange fluid is directed through a second flowpath of the heat exchanger during the first mode of operation. At least some of the heat energy is transferred from the first heat exchange fluid into the second heat exchange fluid. A third heat exchange fluid is directed through a third flowpath of the heat exchanger during the second mode of operation. At least some of the heat energy is transferred from the first heat exchange fluid into the third heat exchange fluid.
F28D 7/00 - Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall
B64D 33/08 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
F28F 27/02 - Control arrangements or safety devices specially adapted for heat-exchange or heat-transfer apparatus for controlling the distribution of heat-exchange media between different channels
H10N 19/00 - Integrated devices, or assemblies of multiple devices, comprising at least one thermoelectric or thermomagnetic element covered by groups
90.
LOCALIZED THICKENING PLY REINFORCEMENT WITHIN CERAMIC MATRIX COMPOSITE AIRFOIL CAVITY
A CMC component for a gas turbine engine includes an airfoil including a leading edge, a trailing edge, a suction sidewall, and a pressure sidewall having a locally-thickened region, the locally-thickened region being a maximum thickness region of the pressure sidewall. The locally-thickened region has a first thickness, a maximum thickness region of the suction sidewall has a second thickness, and the first thickness is greater than the second thickness.
A method of fabricating a fibrous ceramic preform includes impregnating a plurality of individual ceramic tows with a first polymer binder, incorporating the plurality of impregnated ceramic tows into a ceramic fabric, applying a solution comprising a second polymer binder and a second solvent to the ceramic fabric, and incorporating the ceramic fabric into the preform. The first polymer binder is insoluble in the second solvent.
D06M 15/333 - Macromolecular compounds obtained by reactions only involving carbon-to-carbon unsaturated bonds of unsaturated alcohols or esters thereof of vinyl acetatePolyvinylalcohol
D06M 101/16 - Synthetic fibres, other than mineral fibres
92.
APPARATUS AND METHOD FOR PARTIALLY BLADED ROTOR TEST
A rotor module, including: a blade cluster of a plurality of blades secured to a rotor disk of the rotor module; and counter weights secured to the rotor disk, wherein the counter weights cause a center of gravity of the plurality of blades to be zeroed out with respect to an axis “O” of the rotor disk and wherein the counter weights cause a moment of inertia Jx about an X axis of the rotor disk is equal to a moment of inertia Jy about a Y axis of the rotor disk when the rotor module is rotated about the axis “O”.
A counter weight for use with a rotor module, including: a root portion; a main body portion that extends away from the root portion; a pair of side portions each extending from opposite sides of the main body portion; and at least one plate secured to one of the opposite sides of the main body portion in order to adjust a weight of the counter weight.
During a manufacturing method, a workpiece is provided that includes a face surface and a back surface. A first aperture is machined into the workpiece using an energy beam. The first aperture projects partially into the workpiece from the face surface to an end of the first aperture. A waterjet is directed into the first aperture to machine a second aperture into the workpiece. The second aperture extends from the end of the first aperture to the back surface.
B23K 26/382 - Removing material by boring or cutting by boring
B23K 26/146 - Working by laser beam, e.g. welding, cutting or boring using a fluid stream, e.g. a jet of gas, in conjunction with the laser beamNozzles therefor the fluid stream containing a liquid
95.
INSITU BUILD POWDER CHARACTERIZATION FOR POWDER BED FUSION
A build powder characterization system for an additive manufacturing (AM) machine includes a build powder collection system, a build powder analyzer, and a controller. The build powder collection system configured to collect build powder during an AM machine build campaign. The build powder analyzer configured to receive from the build powder collection system the collected build powder, to analyze selected build powder properties, and to communicate to a controller data reflecting properties of the collected build powder. The controller is configured to process data, received from the build powder analyzer, reflecting properties of the collected build powder.
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
A powder bed fusion (PBF) additive manufacturing (AM) machine includes a build plate configured to function as a platform to support one or more parts built using the PBF AM machine. The build plate further includes a thermal history indicator positioned on the build plate, attached to the build plate, or physically integrated with the build plate such that the thermal history indicator experiences the same temperature as the build plate. The thermal history indicator is configured to display a durable visual indication of achievement of a temperature associated with a post-processing step performed after the one or more parts are built on the build platform.
An assembly for a gas turbine engine according to an example of the present disclosure may include at least one rotatable airfoil, at least one vibration sensor operable to detect vibration of the at least one airfoil at one or more rotational frequencies, and a controller operatively coupled to the at least one vibration sensor. The controller may be operable to determine an accretion level associated with accretion of glass on the at least one airfoil in response to comparing vibration of the at least one airfoil at the one or more rotational frequencies to a vibratory pattern associated with accretion of glass. A method of operation is also disclosed.
A gas turbine engine includes first and second CMC vane arc segments arranged about an engine central axis. Each of the CMC vane arc segments has a platform and an airfoil section that extends off of the platform. The platform defines an axially trailing face, an axially leading face that is circumferentially offset from the axially trailing face, and first and second circumferential faces that extend from the axially trailing face to the axially leading face. The first circumferential face of the first CMC vane arc segment interfaces with the second circumferential face of the second CMC vane arc segment. There is at least one friction damper between the first circumferential face of the first CMC vane arc segment and the second circumferential face of the second CMC vane arc segment.
A method is disclosed for a design geometry of a ceramic matrix composite (CMC) vane arc segment that has a platform and an airfoil section that extends off of the platform. The platform defines an axially trailing face, an axially leading face that is circumferentially offset from the axially trailing face, and first and second circumferential faces. In the method, a modal excitation response of the platform is determined based upon an external engine operation excitation frequency. If the modal excitation response is greater than a target modal excitation response, the design geometry of the platform is adjusted to be detuned with the external engine operation excitation frequency by reducing an overhang distance of the platform in which the modal excitation response at the external engine operation excitation frequency is equal to or lower than the target modal excitation response.
A vane assembly includes a vane having an outer platform with mount structure and an airfoil extending inwardly from the outer platform and formed of ceramic matrix composites (“CMC”). The vane having an internal cooling chamber. A metal baffle is received within the internal cooling chamber and extends beyond the outer platform. A bathtub seal has a central portion receiving the metal baffle. The metal baffle has a central chamber and cooling air holes to deliver cooling air into the internal cooling chamber. The bathtub seal having an inner wall and an outer wall outward of the metal baffle. The inner wall, the outer wall and a base wall define a bathtub seal chamber. The base wall is in contact with the outer platform. The vane is connected to the outer wall. The static structure is received within the bathtub seal chamber. A connection is configured to deliver pressurized air to the bathtub seal chamber. A gas turbine engine is also disclosed.
