An aircraft propulsion includes a core engine section that defines a core flow path where an inlet airflow is compressed, mixed with fuel, and ignited to generate an exhaust gas flow, an inner nacelle assembly that surrounds the core engine section, an outer nacelle assembly that is spaced radially apart from the inner nacelle assembly, and a strut heat exchanger that extends radially between the inner nacelle assembly and the outer nacelle assembly, within the strut heat exchanger a portion of the exhaust gas flow is placed in thermal communication with a second flow for transferring thermal energy.
F02C 7/10 - Chauffage de l'air d'alimentation avant la combustion, p. ex. par les gaz d'échappement au moyen d'échangeurs de récupération de chaleur
F28D 7/16 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation les canalisations étant espacées parallèlement
2.
ACCESSIBLE DEBRIS SEPARATOR FOR HIGH PRESSURE TURBINE OUTSIDE DIAMETER FED STATIC COMPONENTS
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 7/05 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes
F01D 11/10 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un fluide d'obturation, p. ex. de la vapeur
An electrically boosted cooling air system including a high pressure compressor; at least one cooling air supply line coupling the high pressure compressor with at least one of a high pressure turbine and a low pressure turbine; and an electrically powered centrifugal compressor fluidly coupled with the at least one cooling air supply line; wherein the electrically powered centrifugal compressor is configured to provide at least one of a pressure increase and a flow rate increase for cooling air supplied from the high pressure compressor.
A combined gas turbine engine vane and blade outer air seal assembly includes a vane having an airfoil extending from the leading edge to a trailing edge, and has an outer platform. The outer platform has a cooling channel that extends into the airfoil to receive cooling air. The outer platform extends to an integral blade outer air seal to be positioned radially outwardly of a turbine blade in a gas turbine engine. At least a portion of the vane and the blade outer air seal are formed of ceramic matrix composite materials. A gas turbine engine is also disclosed.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
5.
Component mounting and drive in a geared turbofan architecture
A fan drive gear system for a turbine engine includes a sun gear that is configured to be driven by an engine shaft rotatable about an axis. A plurality of intermediate gears are coupled to the sun gear, a ring gear is coupled to the plurality of intermediate gears, and a carrier supports rotation of the plurality of intermediate gears. The carrier includes a gear portion and an accessory component that is coupled to and driven by the gear portion of the carrier.
A blade outer air includes a center web having a radially inner face and a radially outer face, at least one mounting arm extending from the radially outer face, and a graded coating disposed on the radially inner face. The graded coating has an abradable component in a first region and a protective component in a second region. The abradable component has a higher porosity than the protective component. A gas turbine engine and a method of protecting a blade outer air seal are also disclosed.
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p. ex. un élément d'usure, déformable ou contraint de façon élastique
7.
Component mounting and drive in a geared turbofan architecture
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. The sun gear includes a first gear portion and a second gear portion. A plurality of intermediate gears are coupled to the first gear portion of the sun gear, a ring gear is coupled to the plurality of intermediate gears, a carrier supports rotation of the plurality of intermediate gears, and an accessory component is coupled to the second gear portion of the sun gear.
An assembly is provided for an engine. This engine assembly includes an engine structure and a fuel injector. The engine structure includes an injector receptacle and a fuel supply passage. The injector receptacle extends longitudinally through the engine structure along a centerline. The fuel supply passage extends laterally within the engine structure to the injector receptacle. The fuel injector is mated with the injector receptacle. The fuel injector includes a nozzle passage and a nozzle outlet. The nozzle passage spirals about the centerline within the fuel injector towards the nozzle outlet. The nozzle passage fluidly couples the nozzle outlet to the fuel supply passage.
A method includes identifying at least one region-of-interest (ROI) of an article that is to be additively manufactured by powder bed fusion, determining a thermal profile of the ROI, where the thermal profile includes at least a maximum temperature and a cooling rate, determining a geometry of an insulation layer upon which at least one process-equivalent test specimen (PETS) is to be additively manufactured by powder bed fusion such that a thermal profile of the at least one PETS replicates the thermal profile of the at least one ROI, and fabricating the at least one PETS in accordance with the thermal profile of the PETS by using the determined geometry of the insulation layer such that the at least one PETS and the at least one ROI are metallurgically equivalent.
An airfoil includes a platform and an airfoil section. The airfoil section and the platform are formed of a ceramic matrix composite (CMC) having fiber plies and a ceramic matrix, the fiber plies include first and second groups of core fiber plies that respectively define first and second radial tubes that circumscribe first and second internal cavities in the airfoil section. Skin fiber plies define an exterior of the airfoil section, wrap around the first and second groups of core fiber plies, and flare outwardly through a fillet into the platform. Platform fiber plies extend in the platform adjacent the skin fiber plies. There is at least one closed-circuit cooling passage defined in the fiber plies. The cooling passage initiates from the first internal cavity, extends within the platform, and ends at the second internal cavity.
An airfoil includes a platform and an airfoil section that are formed of a ceramic matrix composite that includes core fiber plies, skin fiber plies, platform fiber plies, and a cooling passage that extends through selected ones of the plies to provide a cooling circuit through the airfoil section and into the platform.
A fibrous ceramic preform includes a first surface, a second surface opposite the first surface, at least a first thickness defined between the first surface and the second surface, a first zone having a first plurality of z-channels, and a second zone having a second plurality of z-channels. The first plurality of z-channels are different from the second plurality of z-channels.
A gas turbine engine combustor includes a unitary, monolithic combustor body and a unitary, monolithic combustor body boss formed in the unitary, monolithic combustor body. The unitary, monolithic combustor body includes boss threads and a visual boss clocking indicator. The unitary, monolithic combustor body boss is configured to receive, with the boss threads, a fuel injector bolt. The fuel injector bolt includes a bolt head having a visual fuel injector clocking indicator. a threaded body that includes fuel injector threads, one or more fuel intakes, a fuel injector channel, and an angled face that includes a fuel orifice. The fuel injector threads include a fastener arresting face.
An abrasive flow machining (AFM) system includes a first abrasive media reservoir, means for connecting the first abrasive media reservoir to an additive manufacturing (AM) system build plate, and means for directing abrasive media from the first abrasive media reservoir through the plurality of abrasive media flow channels in the AM system build plate into a plurality of flow channels in the structure. The AM system build plate includes a unitized additive structure built on the AM system build plate during an AM system build campaign and a plurality of abrasive media flow channels extending entirely through a thickness of the AM system build plate to facilitate an in-process surface finishing operation.
A system includes at least one drone that is equipped with an imaging device and a light source. At least one processor is configured to: operate the at least one drone to fly into a gas turbine engine to a first position with respect to a component in the gas turbine engine, operate the imaging device and the light source to take an image a target surface of the component from the first position, identify whether the image includes an obstruction blocking a portion of the target surface from view of the imaging device, in response to identifying the obstruction, operate the at least one drone to fly to a second position from which there is a line-of-sight to the target surface without the obstruction, and operate the imaging device and the light source to obtain an unobstructed image of the target surface from the second position.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
A gas turbine engine includes a fan duct, an engine casing, and a gaspath casing. A thermal imaging device is integrated with the gas turbine engine. The thermal imaging device includes: a main body comprising processing circuitry, where the main body is mountable to the fan duct or a mounting structure separate from the fan duct. The thermal imaging device includes a probe mounted to the engine casing. The probe includes one or more sensors configured to detect thermal energy, and the probe is configured to transmit an optical signal based on the detected thermal energy, to an aperture of the main body, over an air gap between the probe and the main body. The processing circuitry is configured to provide temperature information in response to processing the optical signal.
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
17.
METHOD FOR PRODUCING A ROTOR BLADE WITH MOMENT WEIGHT DATA
A method of producing a fan blade with moment weight data is provided that includes: providing a primary master fan blade (PMFB) having a recorded PMFB moment weight value and a recorded PMFB weight value; using a moment weight machine to determine a PMFB moment weight offset value for the primary master fan blade; providing a secondary master fan blade (SMFB); producing a SMFB statistical moment weight value using the moment weight machine; producing SMFB calibration values for the secondary master fan blade using the PMFB moment weight offset value and the SMFB statistical moment weight value; and using the SMFB calibration values and the secondary master fan blade to produce a fan blade with moment weight data.
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.
B29C 51/00 - Façonnage par thermoformage, p. ex. façonnage de feuilles dans des moules en deux parties ou par emboutissage profondAppareils à cet effet
B29C 51/26 - Éléments constitutifs, détails ou accessoiresOpérations auxiliaires
B29K 29/00 - Utilisation de poly(alcool de vinyle), poly(éthers de vinyle), poly(aldéhydes de vinyle), poly(cétones de vinyle) ou poly(cétals de vinyle) comme matière de moulage
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 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.
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 - Analyse de fiabilité ou optimisation de fiabilitéAnalyse de défaillance, p. ex. performance dans le pire scénario, analyse du mode de défaillance et de ses effets [FMEA]
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 - Coulée dans, sur, ou autour d'objets formant partie intégrante du produit final pour fabriquer des moulages composites à partir de métaux différents, p. ex. pour fabriquer des cylindres de laminoirs
B22D 21/00 - Coulée de métaux non ferreux ou de composés métalliques, dans la mesure où leurs propriétés métallurgiques affectent le procédé de couléeUtilisation de compositions appropriées
B23K 9/095 - Surveillance ou commande automatique des paramètres de soudage
B23K 11/00 - Soudage par résistanceSectionnement par chauffage par résistance
B23K 13/01 - Soudage par chauffage au moyen d'un courant haute fréquence par chauffage par induction
B23K 101/00 - Objets fabriqués par brasage, soudage ou découpage
B23P 15/04 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en plusieurs pièces
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
24.
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 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux en utilisant la tomographie, p. ex. la tomographie informatisée
G01B 11/06 - Dispositions pour la mesure caractérisées par l'utilisation de techniques optiques pour mesurer la longueur, la largeur ou l'épaisseur pour mesurer l'épaisseur
G01N 17/00 - Recherche de la résistance des matériaux aux intempéries, à la corrosion ou à la lumière
G01N 21/17 - Systèmes dans lesquels la lumière incidente est modifiée suivant les propriétés du matériau examiné
G01N 21/88 - Recherche de la présence de criques, de défauts ou de souillures
25.
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.
B29B 11/16 - Fabrication de préformes caractérisées par la structure ou la composition comprenant des charges ou des agents de renforcement
B29K 229/00 - Utilisation de poly(alcool de vinyle), poly(éthers de vinyle), poly(aldéhydes de vinyle), poly(cétones de vinyle) ou poly(cétals de vinyle) comme matière de renforcement
C04B 35/657 - Procédés comportant une étape de fusion pour la fabrication de réfractaires
C04B 35/80 - Fibres, filaments, "whiskers", paillettes ou analogues
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
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.
G06V 10/82 - Dispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant les réseaux neuronaux
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 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 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 - Dispositifs amorçant la mise en œuvre actionnés automatiquement
B60L 50/60 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant de la puissance de propulsion fournie par des batteries ou des piles à combustible utilisant de l'énergie fournie par des batteries
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur ou l'énergie de ressorts
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gazCommande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
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.
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 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.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F01D 17/08 - Aménagement des éléments sensibles sensibles aux conditions de fonctionnement du fluide énergétique, p. ex. à la pression
36.
Turbine blade with boomerang shaped wall cooling passages
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.
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 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes comportant des dispositifs séparateurs de poussière
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
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 - Commande de la position ou du cap par référence à un système à deux dimensions
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
G05D 1/00 - Commande de la position, du cap, de l'altitude ou de l'attitude des véhicules terrestres, aquatiques, aériens ou spatiaux, p. ex. utilisant des pilotes automatiques
39.
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 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion
F16H 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
F16H 57/08 - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
40.
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 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes comportant des dispositifs séparateurs de poussière
41.
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 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes comportant des dispositifs séparateurs de poussière
F01D 25/32 - Recueil de l'eau de condensationDrainage
42.
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.
C09D 1/00 - Compositions de revêtement, p. ex. peintures, vernis ou vernis-laques, à base de substances inorganiques
C09D 5/00 - Compositions de revêtement, p. ex. peintures, vernis ou vernis-laques, caractérisées par leur nature physique ou par les effets produitsApprêts en pâte
F01D 25/00 - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
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.
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 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F16N 29/00 - Dispositifs particuliers dans les installations ou les systèmes de lubrification indiquant ou détectant des conditions indésirablesUtilisation des dispositifs sensibles à ces conditions dans les installations ou les systèmes de lubrification
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.
F02C 7/28 - Agencement des dispositifs d'étanchéité
F16J 15/34 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par bague glissante pressée contre la face plus ou moins radiale d'une des deux parties
48.
NEURAL NETWORK BASED FAST MODEL PREDICTIVE CONTROL FOR POWER REGULATION
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 - Procédés ou agencements de circuits pour surveiller ou commander des batteries ou des piles à combustible, spécialement adaptés pour des véhicules électriques pour la surveillance et la commande des batteries
G05B 13/02 - Systèmes de commande adaptatifs, c.-à-d. systèmes se réglant eux-mêmes automatiquement pour obtenir un rendement optimal suivant un critère prédéterminé électriques
G05B 13/04 - Systèmes de commande adaptatifs, c.-à-d. systèmes se réglant eux-mêmes automatiquement pour obtenir un rendement optimal suivant un critère prédéterminé électriques impliquant l'usage de modèles ou de simulateurs
H02J 1/08 - Systèmes à trois filsSystèmes ayant plus de trois fils
49.
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.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
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 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 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur avec plusieurs rotors raccordés par transmission de puissance
F02C 3/10 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur avec une autre turbine entraînant un arbre de sortie mais n'entraînant pas le compresseur
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
53.
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 - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
C23C 18/12 - Revêtement chimique par décomposition soit de composés liquides, soit de solutions des composés constituant le revêtement, ne laissant pas de produits de réaction du matériau de la surface dans le revêtementDépôt par contact par décomposition thermique caractérisée par le dépôt sur des matériaux inorganiques, autres que des matériaux métalliques
C23C 24/08 - Revêtement à partir de poudres inorganiques en utilisant la chaleur ou une pression et la chaleur
C23C 26/00 - Revêtements non prévus par les groupes
C23C 28/00 - Revêtement pour obtenir au moins deux couches superposées, soit par des procédés non prévus dans un seul des groupes principaux , soit par des combinaisons de procédés prévus dans les sous-classes et
C23C 28/04 - Revêtements uniquement de matériaux inorganiques non métalliques
C23C 30/00 - Revêtement avec des matériaux métalliques, caractérisé uniquement par la composition du matériau métallique, c.-à-d. non caractérisé par le procédé de revêtement
C25D 13/18 - Revêtement électrophorétique caractérisé par le procédé utilisant un courant modulé, pulsé ou inversé
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 - Aménagement, montage ou entraînement des auxiliaires
F16H 3/72 - Transmissions à engrenages pour transmettre un mouvement rotatif à rapport de vitesse variable ou pour inverser le mouvement rotatif utilisant des engrenages à mouvement orbital avec entraînement secondaire, p. ex. un moteur régulateur, pour faire varier la vitesse d'une manière continue
F16H 48/10 - Transmissions différentielles avec des engrenages à mouvement orbital avec des engrenages orbitaux droits
57.
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 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel du fluide de travail
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 3/30 - Addition d'eau, de vapeur ou d'autres fluides aux composants combustibles ou au fluide de travail avant l'échappement de la turbine
58.
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.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
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 - Techniques d’interaction fondées sur les interfaces utilisateur graphiques [GUI] pour la commande de fonctions ou d’opérations spécifiques, p. ex. sélection ou transformation d’un objet, d’une image ou d’un élément de texte affiché, détermination d’une valeur de paramètre ou sélection d’une plage de valeurs pour la transformation d’images, p. ex. glissement, rotation, agrandissement ou changement de couleur
G06F 3/0482 - Interaction avec des listes d’éléments sélectionnables, p. ex. des menus
G06T 17/20 - Description filaire, p. ex. polygonalisation ou 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 - CAO géométrique caractérisée par des moyens d’entrée spécialement adaptés à la CAO, p. ex. interfaces utilisateur graphiques [UIG] spécialement adaptées à la CAO
G06T 17/20 - Description filaire, p. ex. polygonalisation ou tessellation
65.
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.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
66.
CORE COMPARTMENT VENT DURING ENGINE SHUTDOWN TO REDUCED BOWED ROTOR START
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 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion
F02C 3/14 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail caractérisés par l'aménagement de la chambre de combustion dans l'ensemble
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 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.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
70.
COATING PROTECTION FOR STATOR VANES AND METHODS OF PROTECTION THEREOF
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 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
C23C 24/10 - Revêtement à partir de poudres inorganiques en utilisant la chaleur ou une pression et la chaleur avec formation d'une phase liquide intermédiaire dans la couche
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
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 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 - CAO géométrique caractérisée par des moyens d’entrée spécialement adaptés à la CAO, p. ex. interfaces utilisateur graphiques [UIG] spécialement adaptées à la CAO
G06F 3/0483 - Interaction avec des environnements structurés en pages, p. ex. métaphore livresque
G06F 3/04855 - Interaction avec des barres de défilement
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 - CAO géométrique caractérisée par des moyens d’entrée spécialement adaptés à la CAO, p. ex. interfaces utilisateur graphiques [UIG] spécialement adaptées à la CAO
G06F 3/0482 - Interaction avec des listes d’éléments sélectionnables, p. ex. des menus
G06T 11/60 - Édition de figures et de texteCombinaison de figures ou de texte
G06T 17/20 - Description filaire, p. ex. polygonalisation ou 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 - Commande numérique [CN], c.-à-d. machines fonctionnant automatiquement, en particulier machines-outils, p. ex. dans un milieu de fabrication industriel, afin d'effectuer un positionnement, un mouvement ou des actions coordonnées au moyen de données d'un programme sous forme numérique caractérisée par le déroulement du programme, c.-à-d. le déroulement d'un programme de pièce ou le déroulement d'une fonction machine, p. ex. choix d'un 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 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.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe
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 - Procédés de mise en formeTraitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques
C04B 35/565 - Produits céramiques mis en forme, caractérisés par leur compositionCompositions céramiquesTraitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques à base de non oxydes à base de carbures à base de carbure de silicium
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 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 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur avec plusieurs rotors raccordés par transmission de puissance
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux comprenant une soufflante avant
86.
GRID ARRAYED MICROTUBE HEAT EXCHANGER WITH MIDSPAN SUPPORT COMPONENTS
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 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation les canalisations étant espacées parallèlement
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 16/455 - Revêtement chimique par décomposition de composés gazeux, ne laissant pas de produits de réaction du matériau de la surface dans le revêtement, c.-à-d. procédés de dépôt chimique en phase vapeur [CVD] caractérisé par le procédé de revêtement caractérisé par le procédé utilisé pour introduire des gaz dans la chambre de réaction ou pour modifier les écoulements de gaz dans la chambre de réaction
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 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 - Addition d'eau, de vapeur ou d'autres fluides aux composants combustibles ou au fluide de travail avant l'échappement de la turbine
F02C 3/34 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail avec recyclage d'une partie du fluide de travail, c.-à-d. cycles semi-fermés comportant des produits de combustion dans la partie fermée du cycle
91.
TOWER SHAFT DRIVEN PLANETARY FAN DRIVE GEAR LUBRICATION SYSTEM
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.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
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 - Ensembles fonctionnels caractérisés par plus d'une machine motrice fournissant de l'énergie à l'extérieur de l'ensemble, ces machines motrices étant entraînées par des fluides différents les cycles de ces machines motrices étant couplés thermiquement la chaleur de combustion provenant de l'un des cycles chauffant le fluide dans un autre cycle le fluide à la sortie de l'un des cycles chauffant le fluide dans un autre cycle
F01K 3/26 - Ensembles fonctionnels caractérisés par l'emploi d'accumulateurs de vapeur ou de chaleur ou bien de réchauffeurs intermédiaires de vapeur comportant des réchauffeurs avec chauffage par la vapeur
F01K 7/16 - Ensembles fonctionnels de machines à vapeur caractérisés par l'emploi de types particuliers de machines motricesEnsembles fonctionnels ou machines motrices caractérisés par un circuit de vapeur, un cycle de fonctionnement ou des phases particuliersDispositifs de commande spécialement adaptés à ces systèmes, cycles ou phasesUtilisation de la vapeur soutirée ou de la vapeur d'évacuation pour le réchauffage de l'eau d'alimentation les machines motrices étant uniquement du type turbine
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 16/9538 - Présentation des résultats des requêtes
G06F 16/25 - Systèmes d’intégration ou d’interfaçage impliquant les systèmes de gestion de bases de données
G06F 16/27 - Réplication, distribution ou synchronisation de données entre bases de données ou dans un système de bases de données distribuéesArchitectures de systèmes de bases de données distribuées à cet effet
94.
Partial exhaust gas condensation with inverse Brayton control
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 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.
G01H 9/00 - Mesure des vibrations mécaniques ou des ondes ultrasonores, sonores ou infrasonores en utilisant des moyens sensibles aux radiations, p. ex. des moyens optiques
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
96.
Coating Repair for Ceramic Matrix Composite (CMC) Substrates
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.
F01D 5/00 - AubesOrganes de support des aubesDispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
97.
PROPULSOR BLADE IMAGING ASSEMBLY FOR AN AIRCRAFT PROPULSION SYSTEM
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.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
H04N 7/18 - Systèmes de télévision en circuit fermé [CCTV], c.-à-d. systèmes dans lesquels le signal vidéo n'est pas diffusé
H04N 23/90 - Agencement de caméras ou de modules de caméras, p. ex. de plusieurs caméras dans des studios de télévision ou des stades de sport
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.
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.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
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 - Silencieux ou dispositifs d'échappement comportant des moyens pour purifier, rendre inoffensifs ou traiter les gaz d'échappement pour rendre les gaz d'échappement inoffensifs par conversion thermique ou catalytique des composants nocifs des gaz d'échappement caractérisés par les méthodes d'opérationCommande
F01N 3/24 - Silencieux ou dispositifs d'échappement comportant des moyens pour purifier, rendre inoffensifs ou traiter les gaz d'échappement pour rendre les gaz d'échappement inoffensifs par conversion thermique ou catalytique des composants nocifs des gaz d'échappement caractérisés par les aspects de structure de l'appareillage de conversion
F02C 7/143 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel du fluide de travail avant ou entre les étages du compresseur
