A gas turbine engine, comprising: a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner that at least partially defines a combustion chamber; and a fuel nozzle assembly, comprising: a liquid fuel supply to supply a liquid fuel; a hydrogen fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle body fluidly coupled with the liquid fuel supply and the hydrogen fuel supply to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber.
A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; and a fuel injector assembly comprising a mixing tube having a mixing tube body defining a mixing channel. The mixing tube body can have a set of fuel passages terminating in fuel orifices fluidly coupled to the mixing channel. A set air flow passages terminating in air outlets can be fluidly coupled to the mixing channel. At least some of the air outlets circumscribe a corresponding fuel orifice.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
F23R 3/02 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la configuration du flux d'air ou du flux de gaz
A turbine engine, comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section further comprising a combustor liner at least partially defining a combustion chamber, and a gaseous fuel nozzle assembly. The gaseous fuel nozzle assembly comprises a set of mixing tubes, a set of fuel jets and an air passage, the air passage including a set of air passage outlets, with each air passage of the set of air passage outlets tangentially, fluidly coupled to a corresponding mixing tube air inlet.
A fuel injector for a gas turbine engine includes a compression section, combustion section, and turbine section in serial flow arrangement. The fuel injector arranged in the combustion section and includes an inner nozzle defining a fuel injector axis and an outer nozzle in annular arrangement about the inner nozzle. The outer nozzle at least partially defines a mixing region. An inner air passage and an outer air passage exhaust to the mixing region. A fuel body positions between the inner air passage and the outer air passage and has an outer fuel passage exhausting to the mixing region.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency are disclosed. An apparatus to support an aircraft engine includes a first thrust link including a forward end and an aft end, the forward end of the first thrust link coupled to the aircraft engine, a second thrust link including a forward end and an aft end, the forward end of the second thrust link coupled to the aircraft engine, and a damper coupled to the aft end of the first thrust link and to the aft end of the second thrust link.
A high entropy alloy-based composition is provided that has the formula: (M1aM2bM3cM4dM5eM6f)CrAlY1-x-zZrxMoz where: each of M1, M2, M3, M4, M5, and M6 is a different alloying element selected from the group consisting of Ni, Co, Fe, Si, Mn, and Cu such that none of M1, M2, M3, M4, M5, and M6 are the same alloying element; 0.05≤a≤0.35; 0.05≤b≤0.35; 0.05≤c≤0.35; 0.05≤d≤0.35; 0.05≤e≤0.35; 0≤f≤0.35; a+b+c+d+e+f=1; 0≤x≤1; 0≤z≤1; and 0≤x+z≤1.
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
A composite airfoil assembly and method of forming includes a spar assembly with a spar core and a spar fiber layer at least partially surrounding the spar core, as well as a root assembly with a sleeve assembly carrying a set of spaced wedges defining at least one slot configured to receive a first end of the spar assembly.
A rotor system for a turbine engine. The rotor system includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that rotate. The stator assembly includes a plurality of stator vanes arranged circumferentially about the stator assembly and includes at least one pair of non-uniform gaps between adjacent stator vanes. The plurality of stator vanes includes a first group of stator vanes having a first non-uniform gap between adjacent stator vanes, a second group of stator vanes having a second non-uniform gap between adjacent stator vanes, and a third group of stator vanes having a uniform spacing between adjacent stator vanes. The first non-uniform gap is positioned 180° from the second non-uniform gap. The plurality of rotor blades directs air through the plurality of stator vanes.
A method of assembling a gas turbine assembly includes providing a core gas turbine engine including a high-pressure compressor, a combustor, and a turbine, coupling a. A low-pressure turbine is axially aft from the core gas turbine engine, coupling a fan assembly is axially forward from the core gas turbine engine, and coupling a booster compressor is coupled to the low-pressure turbine such that the booster compressor and the low-pressure turbine rotate at a first rotational speed, and an epicyclic gearbox is coupled to the low-pressure turbine and the fan assembly such that the fan assembly rotates at a second rotational speed.
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
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
11.
Gas turbine engine, fuel nozzle assembly, and method
A gas turbine engine, comprising a compressor section, combustion section, and turbine section in serial flow arrangement, with the combustion section comprising: a combustor liner at least partially defining a combustion chamber; a wall coupled to the combustor liner; a first fuel supply to supply a first fuel; a gaseous fuel supply to supply a gaseous hydrogen fuel; and a fuel nozzle assembly coupled to the wall and fluidly coupled to the first fuel supply and the gaseous hydrogen fuel supply, the fuel nozzle assembly comprising: a main mixer; and a fuel nozzle comprising an outer wall defining a pilot channel and an outer wall fuel orifice to emit at least one of the first fuel or the gaseous fuel radially outward into the main mixer; and a secondary mixer disposed in the pilot channel.
F23R 3/36 - Alimentation en combustibles différents
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 7/22 - Systèmes d'alimentation en combustible
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F23D 17/00 - Brûleurs pour la combustion simultanée ou alternative de combustibles gazeux, liquides ou pulvérulents
F23R 3/14 - Aménagements de l'entrée d'air pour l'air primaire créant un tourbillon au moyen d'ailettes de tourbillonnement
12.
COMBUSTOR SIZE RATING FOR A GAS TURBINE ENGINE USING HYDROGEN FUEL
A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.
An aircraft defining a vertical direction, a longitudinal direction, and a transverse direction is provided. The aircraft includes: a fuselage; a propulsion system comprising a power source and a plurality of vertical thrust electric fans driven by the power source; and a wing extending from the fuselage in the transverse direction. The wing includes a support structure that comprises a plurality of first support members and a plurality of second support members. The plurality of first support members extending at least partially between the plurality of second support members. The plurality of vertical thrust electric fans arranged between the plurality of first support members and the plurality of second support members.
An airfoil assembly has an airfoil portion and a spar. The airfoil portion has an outer wall. The outer wall extends between a root and a tip in a spanwise direction, and between a leading edge and a trailing edge in a chordwise direction. The outer wall defines an interior. The spar has a spar body. The spar is coupled to the airfoil portion. The spar extends into the interior.
A method of decentralized collaborative target searching, target tracking, or both, includes initializing a probability distribution over a probabilistic search area with an a priori probability distribution of target locations, transmitting the probability distribution to a plurality of vehicles, capturing, using a plurality of sensors on the plurality of vehicles, sensor data relating to locations of a target, updating the a priori probability distribution based on the sensor data or based on data observations outside of the plurality of vehicles, or both, to provide an updated probability distribution, determining optimal trajectories for the plurality of vehicles using the updated probability distribution and using an ergodic principle to define optimality and based on each vehicle of the plurality of vehicles calculating its own optimal trajectory to enable decentralized calculations, and detecting or tracking, or both, the target based on the determining of the optimal trajectories.
G06N 7/01 - Modèles graphiques probabilistes, p. ex. réseaux probabilistes
H04W 4/46 - Services spécialement adaptés à des environnements, à des situations ou à des fins spécifiques pour les véhicules, p. ex. communication véhicule-piétons pour la communication de véhicule à véhicule
Resuscitation apparatus, namely, a component of an infant warmer, namely, an advanced configuration solution set to support resuscitation efforts featuring 3-lead electrocardiograph monitoring, blood oxygen saturation monitoring, integrated resuscitation, scale and Apgar time.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
B64D 27/12 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur des ailes ou fixés à celles-ci
B64C 11/48 - Ensembles de plusieurs hélices coaxiales
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
F02C 6/00 - Ensembles fonctionnels multiples de turbines à gazCombinaisons d'ensembles fonctionnels de turbines à gaz avec d'autres appareilsAdaptations d'ensembles fonctionnels de turbines à gaz à des applications particulières
A turbofan engine includes a fan and a fan actuation system (FAS). The fan is coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The FAS is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The FAS is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine includes a fan and a fan actuation system (FAS). The fan is coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The FAS is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The FAS is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine includes a fan and a fan actuation system (FAS). The fan is coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The FAS is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The FAS is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings. The turbofan engine can also include a gearbox efficiency rating or an overall engine efficiency rating.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
A turbo engine for an aircraft includes a gas turbine engine having a combustion section and a fuel system to provide pressurized fuel to the combustion section. The fuel system includes a fuel tank and a boost pump and a main pump driven by an accessory gearbox that is powered by a shaft of the turbo engine. The boost pump and the main pump are arranged in series. The fuel system also includes an auxiliary pump and a controller to operate the fuel system in (1) a first mode in which the boost pump and the main pump produce pressurized fuel for the turbo engine and the auxiliary pump is deactivated, and (2) a second mode in which the auxiliary pump is activated and produces the pressurized fuel for the turbo engine while bypassing the boost pump and the main pump.
A coated component for coke abatement in a gas turbine engine. The coated component includes a substrate, and a catalytic coating. The catalytic coating includes a phase enriched in metal oxide and a phase enriched in noble metal. The metal oxide is of formula AxEyLzMOu where A is one or more alkaline earth elements, x ranges from zero to one, E is one or more alkali metals, y ranges from zero to one, L is one or more lanthanide elements, z ranges from zero to one, M is one or more d-block or p-block elements, O is oxygen, and u ranges from 0.95 to six.
F02C 7/30 - Prévention de la corrosion dans les espaces balayés par les gaz
B01J 23/68 - Argent ou or avec de l'arsenic, de l'antimoine, du bismuth, du vanadium, du niobium, du tantale, du polonium, du chrome, du molybdène, du tungstène, du manganèse, du technétium ou du rhénium
B01J 23/89 - Catalyseurs contenant des métaux, oxydes ou hydroxydes métalliques non prévus dans le groupe du cuivre ou des métaux du groupe du fer combinés à des métaux nobles
B01J 35/00 - Catalyseurs caractérisés par leur forme ou leurs propriétés physiques, en général
B01J 35/30 - Catalyseurs caractérisés par leur forme ou leurs propriétés physiques, en général caractérisés par leurs propriétés physiques
B01J 35/40 - Catalyseurs caractérisés par leur forme ou leurs propriétés physiques, en général caractérisés par leurs dimensions, p. ex. granulométrie
B01J 37/34 - Irradiation ou application d'énergie électrique, magnétique ou ondulatoire, p. ex. d'ondes ultrasonores
B08B 17/02 - Procédés pour empêcher la salissure pour empêcher le dépôt de crasses ou de poussières
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
F23R 3/40 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'emploi de moyens catalytiques
23.
METHOD FOR BONDING CMC FACESHEETS TO CERAMIC CORES
A method for forming a ceramic matrix composite (“CMC”) component includes applying a CMC fiber preform to a ceramic core where the CMC fiber preform is formed from a plurality of CMC fiber tows. An interface coating is formed on at least a portion of the plurality of CMC fiber tows before or after the CMC fiber preform is formed from the plurality of CMC fiber tows. A ceramic matrix is formed on the CMC fiber preform with the CMC fiber preform applied to the ceramic core. The CMC fiber preform and the ceramic matrix form a CMC facesheet. A reaction material is located at or adjacent an interface of the CMC facesheet and the ceramic core. The CMC facesheet and the ceramic core are thermally processed to react the interface coating with the reaction material to form a bonding layer at the interface.
A ceramic matrix composite (“CMC”) component includes a ceramic core comprising one or more interlocking elements. A CMC facesheet is positioned against a surface of the ceramic core and in mechanical engagement with the one or more interlocking elements to secure the CMC facesheet to the ceramic core. At least a portion of CMC facesheet is disposed between at least one interlocking element of the one or more interlocking elements and the surface.
A gas turbine engine including a fan section, a compressor section, combustion section. An airfoil is provided in one of the fan section, the compressor section, or the turbine section, and includes an outer wall defining an interior. A first portion is positioned within the interior and at least one disengagement feature is positioned within the first portion to disengage or separate within the first portion in response to a force exceeding a predetermined threshold.
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 turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
An energy management system and method of operating the energy management, which include estimating an energy demand for flight plans for a fleet of aircraft. The flight plans are received from a flight plan database. The system is configured to determine whether a set of dischargeable energy modules are locatable at a respective location of a subset of the fleet of aircraft based at least in part on a replaceable power source inventory database or the subset of the plurality of flight plans. The system is configured to generate a power source inventory distribution plan allocating a subset of dischargeable energy modules for the subset of the plurality of flight plans for the fleet of aircraft based at least in part on the determination that the set of dischargeable energy modules are locatable at the respective location of the subset of the fleet of aircraft.
H02J 7/00 - Circuits pour la charge ou la dépolarisation des batteries ou pour alimenter des charges par des batteries
B60L 50/50 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant de la puissance de propulsion fournie par des batteries ou des piles à combustible
B60L 58/18 - 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 de plusieurs modules de batterie
B64D 27/355 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles à combustible
B64D 27/357 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles
B64D 27/359 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des condensateurs
A power distribution system for an aircraft, and method of operating includes, including determining, by a supplemental electrical power system, an operating mode of the aircraft, selecting, by the supplemental electrical power system, at least one supplementary power source of a set of at least two supplementary power sources, and based on the operating mode of the aircraft at least one of receiving, by the selected at least one supplementary power source, a first current from the primary power bus, or selectively providing a second current to the primary power bus from the selected at least one supplementary power source.
H02J 1/10 - Fonctionnement de sources à courant continu en parallèle
H02J 7/00 - Circuits pour la charge ou la dépolarisation des batteries ou pour alimenter des charges par des batteries
H02J 7/14 - Circuits pour la charge ou la dépolarisation des batteries ou pour alimenter des charges par des batteries pour la charge de batteries par des générateurs dynamo-électriques entraînés à vitesse variable, p. ex. sur véhicule
H02M 3/335 - Transformation d'une puissance d'entrée en courant continu en une puissance de sortie en courant continu avec transformation intermédiaire en courant alternatif par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrodes de commande pour produire le courant alternatif intermédiaire utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs
30.
TURBINE ENGINE FOR AN AIRCRAFT INCLUDING A CONTRAIL MITIGATION SYSTEM
A turbine engine for an aircraft includes a fuel delivery assembly for a hydrocarbon fuel to flow therethrough, a combustor combusting the fuel to generate combustion gases, and a core air exhaust nozzle exhausting the combustion gases from the turbine engine. The turbine engine also includes a contrail mitigation system having a heater and a fuel precipitate separator. The heater is selectively operable to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel, and the fuel precipitate separator separates the fuel precipitates generated by the heater from the fuel. A controller is coupled to the heater to operate the heater to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel in response to a contrail mitigation input.
A gas turbine engine is provided, comprising: a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order; and a fan section having a fan drivingly coupled to the turbomachine and an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine, the airflow surface defining a plurality of boundary layer openings configured to ingest a boundary layer of the fan airflow over the airflow surface during operation of the gas turbine engine.
A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
F02K 3/065 - 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 des soufflantes avant et arrière
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
An insertion tool with an implement gripping mechanism is provided. The insertion tool includes an elongated section defining an implement path extending from a proximal end to a distal end, a handle coupled to the proximal end of the elongated section, a grip mechanism coupled to the handle and configured to selectively secure an implement inserted into the implement path through the handle, and a grip actuator operable to cause the grip mechanism to secure the implement to the handle.
A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.
A heat exchanger includes an outer body defining a centerline axis and a flow passage extending from an inlet to an outlet of the heat exchanger. A first annular fin is disposed within the flow passage, wherein the first annular fin is concentrically aligned with the centerline axis and defines a first undulating surface. A second annular fin is disposed within the flow passage. The second annular fin is concentrically aligned with the centerline axis and defines a second undulating surface. The second annular fin is radially spaced from the first annular fin to define a flow channel therebetween. The flow passage, the first annular fin, and the second annular fin diverge along the centerline axis downstream from the inlet.
A turbine engine comprising a compression section, combustion section, and turbine section is serial flow arrangement, with the combustion section including an injector for providing a mixture of fuel and air for combustion. The injector includes a body, an inner nozzle provided within the body and defining an injector axis, and an outer nozzle in annular arrangement about the inner nozzle. A first fuel passage fluidly couples to the inner nozzle and a second fuel passage fluidly couples to the outer nozzle. A first set of air conduits are in annular arrangement about the body interior of the outer nozzle and a second set of air conduits are in annular arrangement about the body exterior of the outer nozzle.
An aircraft is provided defining a longitudinal direction, a lateral direction, and a longitudinal centerline extending along the longitudinal direction, the aircraft including: a body; a pair of wings extending outward from the body along the lateral direction, each wing of the pair of wings defining a leading edge; and a propulsion system comprising a first engine and a second engine spaced from one another along the lateral direction, the propulsion system further comprising a first thrust vectoring system operable with the first engine and a second thrust vectoring system operable with the second engine.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion, which can supply a primary fuel supply and a secondary fuel supply. Increasing efficiency and reducing emission require the use of alternative fuels, which combust at higher temperatures or burn at faster burn speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.
A fan assembly for a gas turbine engine includes a fan actuation system, a fan blade hub, a plurality of fan blades, and a counterweight system arranged disconnected from the plurality of fan blades. The counterweight system includes a counterweight hub, and a plurality of counterweight levers each having a counterweight trunnion rotationally connected to the counterweight hub, a cantilever arm with a counterweight connected thereto. The fan blades are rotationally connected to the fan blade hub and to the fan actuation system, and the counterweight levers are rotationally connected to the counterweight hub and to the fan actuation system. The fan actuation system is arranged to correspondingly rotate each of the plurality of fan blades about a respective fan blade pitch change axis, and each of the plurality of counterweight levers about a counterweight lever rotational axis in unison.
Sensor information from a plurality of sensors that monitor a heat exchanger is accessed and then sensed information that corresponds to that sensor information is input into a control circuit configured to output virtual parameters corresponding to the heat exchanger as a function, at least in part, of the sensed information. A particular fault class is determined from amongst a plurality of fault classes as a function, at least in part, of the virtual parameters and at least one task is identified regarding the heat exchanger that corresponds to the particular fault class.
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
B64D 13/08 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacements réservés au fret l'air étant climatisé l'air étant réchauffé ou refroidi
F02C 7/12 - Refroidissement des ensembles fonctionnels
42.
PROPULSION SYSTEM HAVING FLUIDLY SEPARATE LUBRICATION SYSTEMS FOR PROPULSORS AND TURBO-ENGINE
A propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine including a turbo-engine, a first propulsor including a first propulsor shaft, and a first gearbox assembly having a gear assembly. The propulsion system includes a second propulsor that is remote from the turbine engine and includes a second gearbox assembly including a gear assembly. The propulsion system further includes a lubrication system including a first gearbox lubrication system, an engine lubrication system, and a second gearbox lubrication system. The first gearbox lubrication system is disposed within the first gearbox assembly and supplies lubricant to the gear assembly. The engine lubrication system supplies lubricant to the engine bearings. The first gearbox lubrication system is fluidly separate from the engine lubrication system. The second gearbox lubrication system is disposed within the second gearbox assembly and supplies lubricant to the gear assembly.
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
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
B64C 11/48 - Ensembles de plusieurs hélices coaxiales
B64D 35/02 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques
B64D 35/04 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions caractérisée par le fait que la transmission entraîne plusieurs hélices ou rotors
An airfoil assembly for a turbine engine, the airfoil assembly including a platform defining an inner surface and an outer surface, a variable pitch airfoil extending radially from the outer surface of the platform from a root to a tip to define a span length and a mounting structure connected to the platform.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor. The unducted fan propulsors can also include a first VPF parameter within a range of 0.10 to 0.40 and defined as the hub-to-tip radius ratio divided by the fan pressure ratio and/or a second VPF parameter within a range of 1-30 lbf/in2 and defined as the bearing spanwise force divided by the fan area. In certain examples, the unducted fan propulsor further includes a pitch change mechanism and/or a gearbox.
B64D 27/18 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à réaction à l'intérieur des ailes ou fixés à celles-ci
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
46.
AUTOMATED DE-POWDERING OF ADDITIVE MANUFACTURING BUILD
An automated de-powdering system comprises a sleeve alignable with at least a portion of a build chamber and at least one actuator actuatable with respect to a build plate to move at least a portion of an additive manufacturing build out of the build chamber and into the sleeve. At least one support assembly is couplable to the sleeve and includes at least one support member insertable through the sleeve and into the additive manufacturing build. At least one agitation mechanism is couplable to at least one of the at least one support assembly or the sleeve and is actuatable to at least partially convey the powder build material away from at least one object of the additive manufacturing build. The support member is positioned to support the object in an absence of the powder build material around the object.
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p. ex. par frittage ou fusion laser sélectif
B33Y 40/20 - Posttraitement, p. ex. durcissement, revêtement ou polissage
A turbofan engine for an aircraft includes a core cowl, a nacelle assembly positioned radially outward of the core cowl defining a bypass airflow passage between the core cowl and the nacelle assembly where the bypass airflow passage has a fan exit nozzle. The nacelle assembly includes a fan cowl, a transcowl positioned aft of the fan cowl, and a thrust reverser assembly. An actuation assembly is operably connected to at least one of the transcowl or the thrust reverser assembly and is actuatable to move the transcowl aft from a first position where the cascade assembly is covered to a second position where the cascade assembly is uncovered. The actuation assembly is further actuatable to move the transcowl forward from the first position to a third position to reduce an area of the fan exit nozzle.
F02K 1/72 - Inversion du flux de la soufflante utilisant des volets inverseurs de poussée ou des portes montées sur le carter de la soufflante la partie arrière du carter de la soufflante étant mobile pour découvrir des ouvertures d'inversion de poussée dans le carter de la soufflante
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
A turbine engine comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising a combustion chamber, a fuel nozzle assembly comprising a fuel supply passage and having a fuel supply passage outlet fluidly coupled to a combustion chamber, and an optical sensor located in the fuel supply passage and oriented to sense a combustion flame in the combustion chamber.
F23N 5/08 - Systèmes de commande de la combustion utilisant des dispositifs sensibles aux variations thermiques ou à la dilatation thermique d'un agent utilisant des éléments sensibles à la lumière
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
49.
Control Logic for Thrust Link Whiffle-Tree Hinge Positioning for Improved Clearances
Systems and methods for optimizing clearances within an engine include an adjustable coupling configured to couple a thrust link to the aircraft engine, an actuator coupled to the adjustable coupling, where motion produced by the actuator adjusts a hinge point of the adjustable coupling, sensors configured to capture real time flight data, and an electronic control unit. The electronic control unit receives flight data from the sensors, implements a machine learning model trained to predict clearance values within the engine based on the received flight data, predicts, with the machine learning model, the clearance values within the engine based on the received flight data, determines an actuator position based on the clearance values, and causes the actuator to adjust to the determined actuator position.
F01D 11/14 - Régulation ou commande du jeu d'extrémité des aubes, c.-à-d. de la distance entre les extrémités d'aubes du rotor et le corps du stator
F01D 11/22 - Réglage actif du jeu d'extrémité des aubes par actionnement mécanique d'éléments du stator ou du rotor, p. ex. par déplacement de sections d'enveloppe par rapport au rotor
F01D 11/24 - Réglage actif du jeu d'extrémité des aubes par refroidissement ou chauffage sélectifs d'éléments du stator ou du rotor
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
Sensor information from a plurality of sensors that monitor an apparatus is accessed and then sensed information that corresponds to that sensor information is input into a control circuit configured to output virtual parameters corresponding to the apparatus as a function, at least in part, of the sensed information. A particular fault class is determined from amongst a plurality of fault classes as a function, at least in part, of the virtual parameters and at least one task is identified regarding the apparatus that corresponds to the particular fault class.
An airfoil for a gas turbine engine. The airfoil includes an inner portion including a hollow fiber composite, and a surface portion including a solid fiber composite. The surface portion at least partially surrounds the inner portion. The hollow fiber composite of the inner portion can be a woven composite with at least one of a warp fiber or a weft fiber including a hollow carbon fiber. The hollow fiber composite of the inner portion also can be a braided composite with at least one strand including a hollow carbon fiber.
A propulsion system and a method of operating the propulsion system. The propulsion system includes a first turbine engine, a second turbine engine, and a lubrication system. The lubrication system supplies a lubricant to the first turbine engine and the second turbine engine. The lubrication system includes one or more first lubricant sensors that sense information about the lubricant to the first turbine engine and one or more second lubricant sensors that sense information about the lubricant to the second turbine engine. The first controller and the second controller receive the information of the lubricant from the first lubricant sensors and the second lubricant sensors. The first controller controls the lubrication system to supply the lubricant to the second turbine engine based on the sensed information about the lubricant in the second turbine engine received at the first controller when the second turbine engine is shut down.
A propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine, an electric power supply, and a lubrication system. The turbine engine includes a turbo-engine having a low-pressure shaft and one or more engine bearings, a propulsor having a propulsor shaft, a gearbox assembly having a gear assembly, and an electric machine. The propulsor shaft is drivingly coupled to the low-pressure shaft through the gear assembly. The electric machine powers the propulsor when the turbo-engine is shut down. The lubrication system supplies a lubricant to at least one of the one or more engine bearings or the gear assembly. The electric power supply powers the lubrication system when the turbo-engine is shut down.
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
A turbine engine control system for a propulsion system and a method of operating the propulsion system. The propulsion system includes a turbine engine including a turbo-engine and a first propulsor drivingly coupled to the turbo-engine, and a second propulsor that is remote from the turbine engine. The turbine engine control system includes a single throttle lever and a controller that receives an input from the single throttle lever, and controls the turbine engine and the second propulsor based on the input from the single throttle lever.
A turbine engine includes a turbo-engine, a fan, a frame, and a lubrication system. The turbo-engine includes a core air flowpath. The fan is drivingly coupled to the turbo-engine. The frame supports the core air flowpath and includes a plurality of lubricant struts that extend through the core air flowpath. The lubrication system includes a sump having a lubricant therein, a scavenge reservoir, and a lubricant strut flowpath disposed through each of the plurality of lubricant struts. The lubricant strut flowpath is in fluid communication with the sump and the scavenge reservoir. The lubricant strut flowpath of each of the plurality of lubricant struts directs the lubricant from the sump to the scavenge reservoir.
A propulsor assembly for a turbine engine includes a propulsor and a propulsor actuation system. The propulsor has a plurality of propulsor blades. Each of the plurality of propulsor blades is rotatable about a blade pitch axis. The propulsor actuation system includes an actuator for rotating the plurality of propulsor blades and a trunnion mechanism that includes a plurality of trunnion assemblies. Each of the plurality of trunnion assemblies is coupled to a respective one of the plurality of propulsor blades and includes an outer sleeve coupled to a blade spar of the respective propulsor blade, an inner sleeve coupled to the outer sleeve, and an actuation member that engages the inner sleeve. The actuation member is rotatably engageable by the actuator to rotate the respective propulsor blade about the blade pitch axis.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F04D 29/32 - Rotors spécialement adaptés aux fluides compressibles pour pompes à flux axial
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
F04D 19/00 - Pompes à flux axial spécialement adaptées aux fluides compressibles
58.
ADDITIVE MANUFACTURING TESTING APPARATUSES AND METHODS FOR USING SAME IN MEASURING TENSILE AND COMPRESSIVE LOADS
An additive manufacturing testing apparatus including an interchangeable base plate, an interchangeable dolly, and a sensor. The interchangeable base plate includes a transparent section over which a photocurable material is disposed. The interchangeable dolly is positionable between an upper position and a lower position along a vertical axis. The interchangeable dolly is configured to compress the photocurable material between the transparent section and the interchangeable dolly and configured to retract from the transparent section. The sensor is coupled to the interchangeable dolly. The sensor is configured to measure tensile load and compressive load during movement of the interchangeable dolly between the upper position and the lower position along the vertical axis during compression of the photocurable material and retraction of the interchangeable dolly from the transparent section.
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain that is coupled in gear with the low pressure spool. The low pressure geartrain is coupled in gear with the low pressure spool and the high pressure geartrain is coupled in gear with the high pressure spool at a common axial location relative to a centerline axis of the gas turbine engine.
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
60.
GAS TURBINE ENGINE HAVING A MECHANICAL POWER SHARING ARRANGEMENT
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain coupled in gear with the low pressure spool. The gas turbine engine further includes a power transfer device coupling the high pressure geartrain and the low pressure geartrain to transfer power between the high pressure spool and the low pressure spool.
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
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
61.
GAS TURBINE ENGINE HAVING A MECHANICAL POWER SHARING ARRANGEMENT
A gas turbine engine includes a low pressure spool that connects a low pressure compressor to a low pressure turbine. The gas turbine engine further includes a high pressure spool that connects a high pressure compressor to a high pressure turbine. The gas turbine engine further includes a high pressure geartrain that is coupled in gear with the high pressure spool. The gas turbine engine further includes a low pressure geartrain that is coupled in gear with the low pressure spool. The gas turbine engine further includes an accessory gearbox that has an accessory drive gear that is selectively drivingly coupled at least one of the high pressure geartrain and the low pressure geartrain.
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
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
A blended wing aircraft is provided, including: an aircraft engine comprising a lubrication system; a body having a fuselage and a pair of wings extending outward from the fuselage; and an access panel assembly operable with the lubrication system of the aircraft engine, the access panel assembly positioned on or within the body at a location remote from the aircraft engine.
A lubrication system for a turbine engine that includes one or more rotating components. The lubrication system includes one or more tanks that store lubricant, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during stable operating conditions of the lubrication system. The auxiliary lubrication system includes an auxiliary feed line and an auxiliary supply line. The auxiliary lubrication system receives the lubricant from the one or more tanks through the auxiliary feed line. The auxiliary lubrication system supplies the lubricant to the one or more rotating components through the auxiliary supply line when there is a potential lubricant interruption in the lubrication system.
A thermal management system for a propulsion system of an aircraft includes a fluid circuit configured to provide a flow of a compressed fluid from a compressor section of the propulsion system to, in serial flow order, a first turbine, a compressor, a second turbine, a thermal load, and an exhaust sink. A heat exchanger is disposed in a fan stream of the propulsion system at a location along the fluid circuit between the compressor and the second turbine. The fluid circuit extends through and is in thermal communication with the heat exchanger.
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
⨯
D
FT
L
AXIAL
⨯
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given by
N
FB
⨯
D
FT
L
AXIAL
⨯
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F04D 29/32 - Rotors spécialement adaptés aux fluides compressibles pour pompes à flux axial
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
F04D 19/00 - Pompes à flux axial spécialement adaptées aux fluides compressibles
F04D 29/54 - Moyens de guidage du fluide, p. ex. diffuseurs
66.
PULSE ILLUMINATION IMAGING OF A MOVING TARGET ELEMENT
A pulse illumination imaging system is provided. The system includes an image sensor, a light source, and a controller. The image sensor includes a plurality of light sensitive pixel elements that are activatable for a designated exposure time to capture one or more images. The controller is configured to determine an activation time to activate the image sensor; activate the image sensor at the activation time; and activate the light source during the exposure time of the image sensor to produce a pulse having a preconfigured time duration that is less than the exposure time of the image sensor.
H04N 23/74 - Circuits de compensation de la variation de luminosité dans la scène en influençant la luminosité de la scène à l'aide de moyens d'éclairage
H04N 23/68 - Commande des caméras ou des modules de caméras pour une prise de vue stable de la scène, p. ex. en compensant les vibrations du boîtier de l'appareil photo
H04N 23/73 - Circuits de compensation de la variation de luminosité dans la scène en influençant le temps d'exposition
A fuel tank heat rejection system for an aircraft. The fuel tank heat rejection system includes a fuel tank compartment in the aircraft and a fuel tank having an exterior surface and storing fuel therein. The fuel tank is located in the fuel tank compartment. The fuel tank heat rejection system includes one or more air valves that provide fluid communication to the fuel tank compartment. The one or more air valves opening to operably direct cooling air into the fuel tank compartment through the one or more air valves, the cooling air contacting the exterior surface of the fuel tank such that heat from the fuel is rejected from the fuel tank.
A gearbox assembly for a turbine engine. The gearbox assembly includes a gearbox having a plurality of gears, and a scavenge gutter radially outward of the gearbox and mounted to a fan frame flange of a fan frame of the turbine engine, the scavenge gutter being configured to collect lubricant that is ejected by the plurality of gears of the gearbox during rotation of the plurality of gears. An interface between the fan frame flange and the scavenge gutter is located at a first radial distance R1 relative to a longitudinal centerline axis of the turbine engine and the scavenge gutter is located at a second radial distance R2 relative to the longitudinal centerline axis. The second radial distance R2 is greater than the first radial distance R1. The scavenge gutter is segmented and includes a plurality of sector portions joined together to form the scavenge gutter.
A turbine engine includes a fan having a plurality of fan blades, a nacelle that surrounds the fan, and a fan blade sensor system. The fan blade sensor system includes a plurality of fan blade sensors disposed in the nacelle to sense the plurality of fan blades as the plurality of fan blades rotates. The fan blade sensor system also includes a controller that determines a fan speed of the fan based on a rotational time between the fan blades as sensed by the plurality of fan blade sensors.
F01D 7/00 - Rotors à aubes réglables en marcheLeur commande
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
70.
SYSTEMS AND METHODS FOR SECURE ADDITIVE MANUFACTURING
Systems, apparatus, computer-readable medium, and associated methods for secure additive manufacturing are disclosed. An example apparatus includes an inbound one-way data diode to receive, authenticate, and route an inbound file in a first direction within a secure additive manufacturing system, the inbound one-way data diode unable to transmit data out of the secure additive manufacturing system in a second direction. The example apparatus includes an additive manufacturing machine to build a part, the build of the part adjusted by the inbound file when authenticated by the inbound one-way data diode. The example apparatus includes an outbound one-way data diode to authenticate and transmit outbound data in the second direction to an external system outside the secure additive manufacturing system, the outbound one-way data diode unable to transmit data into the secure additive manufacturing system in the first direction.
G06F 21/62 - Protection de l’accès à des données via une plate-forme, p. ex. par clés ou règles de contrôle de l’accès
B29C 64/393 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B33Y 50/02 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
G06F 21/57 - Certification ou préservation de plates-formes informatiques fiables, p. ex. démarrages ou arrêts sécurisés, suivis de version, contrôles de logiciel système, mises à jour sécurisées ou évaluation de vulnérabilité
A gas turbine engine defining an axial direction and a radial direction includes a spinner defining a spinner duct and a spinner inlet to the spinner duct, a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a fan duct, a fan duct inlet to the fan duct, a core duct, and a core inlet to the core duct, and a primary fan driven by the turbomachine, wherein the spinner inlet is upstream of the primary fan.
An example booster splitter includes an inner cylindrical structure; an outer cylindrical structure concentric with the inner cylindrical structure; an annular lip having an arc extending between the inner cylindrical structure and the outer cylindrical structure; a first protrusion at a first circumferential position on the annular lip, the first protrusion protruding axially from the annular lip and extending from the outer cylindrical structure to the inner cylindrical structure along the arc of the annular lip; and a second protrusion at a second circumferential position on the annular lip.
F01D 25/02 - Dispositifs de dégivrage pour machines motrices dans lesquelles se produisent des phénomènes de givrage
B64D 15/00 - Dégivrage ou antigivre des surfaces externes des aéronefs
B64D 15/16 - Dégivrage ou antigivre des surfaces externes des aéronefs par dispositifs mécaniques, p. ex. des gaines ou des bourrelets fixés ou incorporés à la surface et soumis à des pulsations
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
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
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
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
A heat exchanger assembly includes a manifold, a plurality of plates supported by the manifold, a bypass channel in fluid communication with the manifold, and a flow controller fluidly connecting a heated fluid supply to the bypass channel. The flow controller is configured to flow a heated fluid from the heated fluid supply through the bypass channel.
F02K 3/065 - 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 des soufflantes avant et arrière
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
76.
BINDER SOLUTIONS COMPRISING A FUGITIVE METAL PRECURSOR FOR USE IN ADDITIVE MANUFACTURING
A binder solution comprises a fugitive metal precursor, a thermoplastic binder, and a solvent. The fugitive metal precursor may comprise an alkaline earth metal, a transition metal, a post-transition metal, a metalloid, a rare earth metal, or combinations thereof. The fugitive metal precursor may comprise a salt such as carboxylate, nitrate, sulfate, carbonate, formate, chloride, halide, derivatives thereof, and combinations thereof. A method of manufacturing a part includes depositing a layer of particulate material on a working surface, selectively applying a binder solution into the layer of particulate material in a pattern representative of a layer of the part, repeating the steps of depositing and selectively applying to form a plurality of layers of particulate material with the applied binder solution, and curing the applied binder solution in the plurality of layers of particulate material with the applied binder solution to evaporate the solvent and form a green body part.
C09D 129/04 - Alcool polyvinyliqueHomopolymères ou copolymères partiellement hydrolysés d'esters d'alcools non saturés avec des acides carboxyliques saturés
B22F 10/14 - Formation d’un corps vert par projection de liant sur un lit de poudre
B22F 10/64 - Traitement de pièces ou d'articles après leur formation par des moyens thermiques
A turbine engine includes a compression section, combustion section, and turbine section is serial flow arrangement. A supply of gaseous fuel is mixed with a supply of inert gas to form a mixture of fuel and inert gas. The mixture of fuel and inert gas is provided to or intermixed within a fuel injector for combustion within a combustor provided within the combustor section to drive the turbine section.
Truss-braced wing aircraft engine mounting and associated systems are disclosed. An example aircraft includes a fuselage including a central structural section, a wing extending from the fuselage, a truss extending from the central structural section and coupled to the wing, a pylon coupled to the central structural section, and an engine coupled to the pylon, wherein the pylon is positioned at an angle substantially 45 degrees from a horizontal plane and a vertical plane.
A combustor having a plurality of openings to control a vortex driver jet within the combustor. The combustor includes an outer casing and an inner casing extending circumferentially about a longitudinal combustor centerline axis, an outer liner spaced apart from the outer casing to define therebetween an outer flow passage and an inner liner spaced apart from the inner casing to define therebetween an inner flow passage, a dome structure, the outer liner and the inner liner defining a combustion chamber, a plurality of outer openings provided in the outer casing and a plurality of inner openings provided in the inner casing. The plurality of outer openings and inner openings are configured to bleed airflow from or to introduce airflow into the outer flow passage or the inner flow passage, to control the vortex driver jet within the driver openings, to drive a vortex in the combustion chamber.
GE Marmara Technology Center Muhendislik Hizmetleri Ltd (Turquie)
GE Aerospace Poland Sp. z o.o. (Pologne)
Inventeur(s)
Unsal, Arda
Whitener, Geoffrey
Bibler, John David
Yilmaz, Batu
Beyer, Katherine
Pazinski, Adam Tomasz
Abrégé
An engine includes a compressor including an inner casing and an outer casing where the inner casing defines a primary flow path for a primary airflow through the compressor. The inner casing and the outer casing define a bleed air cavity therebetween. The inner casing at least partially defines a bleed air channel extending circumferentially about the inner casing to direct a bleed airflow from the primary airflow into the bleed air cavity. One or more flow control devices located circumferentially about the compressor to actively or passively circumferentially balance a flow of the bleed airflow into the bleed air cavity or within the bleed air cavity, wherein the one or more flow control devices are disposed at least partially within the bleed air channel, form at least part of the bleed air channel, or extend axially aft from the bleed air channel.
General Electric Company Polska sp. z o.o. (Pologne)
Inventeur(s)
Li, Cong
Yi, Xuan
Younsi, Karim
Huang, Shenyan
Xiong, Han
Hanczewski, Pawel Piotr
Klausen, Michael
Abrégé
An electrical cable assembly includes a cable portion and a connector housing portion coupled to the cable portion. The cable portion includes a first electrical conductor circumferentially surrounded by a first set of layers. The first set of layers includes a first semiconductive layer, a first insulative layer, a second semiconductive layer, a magnetic layer, a second insulative layer, and an electrically conductive. The connector housing portion includes a connector housing, a printed circuit board disposed within the connector housing; and a set of capacitors mounted to the printed circuit board and coupled to the magnetic layer.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24 and given by
N
FB
×
D
FT
L
AXIAL
×
(
R
TB
N
FB
)
.
NFB is a number of the fan blades, DFT is a fan tip diameter of the fan blades, RTB is a thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length from a fan hub tip to the fan bearings.
F01D 7/02 - Rotors à aubes réglables en marcheLeur commande ayant un réglage sensible à la vitesse
F01D 17/26 - Dispositifs utilisant des éléments sensibles ou des organes de commande terminaux ou les organes de liaison entre les deux, p. ex. commande assistée l'énergie de fonctionnement ou de puissance assistée étant essentiellement non mécanique à fluide, p. ex. hydraulique
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
F03D 7/02 - Commande des mécanismes moteurs à vent les mécanismes moteurs à vent ayant l'axe de rotation sensiblement parallèle au flux d'air pénétrant dans le rotor
F03D 17/00 - Surveillance ou test de mécanismes moteurs à vent, p. ex. diagnostics
F03D 80/70 - Dispositions de roulement ou de graissage
83.
PROPULSION SYSTEM INCLUDING AN ELECTRIC MACHINE FOR STARTING A GAS TURBINE ENGINE
A gas turbine engine includes a turbomachine comprising a low pressure (LP) spool and a high pressure (HP) spool that rotate about a central axis, an electric motor mechanically coupled to the LP spool for selectively rotating the LP spool, a starter assembly mechanically coupled to the HIP spool for selectively rotating the HP spool, and a controller in operative communication with the electric motor and the starter assembly, the controller being configured to operate the electric motor to rotate the LP spool and operate the starter assembly to rotate the HIP spool during engine startup.
F02C 7/268 - Entraînement du rotor pour le démarrage
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F02C 6/00 - Ensembles fonctionnels multiples de turbines à gazCombinaisons d'ensembles fonctionnels de turbines à gaz avec d'autres appareilsAdaptations d'ensembles fonctionnels de turbines à gaz à des applications particulières
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
A turbine component comprised of a titanium alloy that has been modified from Ti-64 is provided. The modification preserves the desired properties of Ti-64 (e.g., relatively isotropic properties, a relatively low density, tolerance to FOD, repairability, and low cost) while improving the thick section strength, HCF capability, creep strength, and low deformation following FOD to approach those beneficial aspects of Ti-17 and Ti-6246. Methods of forming such turbine components are also provided.
Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.
F02C 6/08 - 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 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p. ex. ensembles fonctionnels de chauffage à turbine à gaz
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
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. Increasing efficiency and meeting emission needs can be met with the use of alternative fuels, which combust at higher temperatures or higher speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.
A combustor liner for a combustor of a gas turbine includes an outer liner having a plurality of outer liner segments, and an inner liner having a plurality of inner liner segments. Each segment of the inner liner and the outer liner includes at least one slotted dilution opening therethrough extending in a circumferential direction, and each slotted dilution opening includes a deflector wall extending radially from the respective liner into a dilution zone of a combustion chamber between the outer liner and the inner liner. The at least one slotted dilution opening may be a curved slot (either concave or convex) dilution opening, and the curved slot dilution openings for each segment of the outer liner and the inner liner may be connected so as to provide a wavy slotted dilution opening extending annularly through the outer liner and the inner liner.
An apparatus for maintaining a gas turbine engine having at least one port comprises a tool having an end effector to effect maintaining the gas turbine engine, the tool being configured to temporarily enter and exit the gas turbine engine via the at least one port and having a first portion and a second portion that are separated by at least a first area of articulation. A first inertial measurement unit is affixed with respect to that first portion and a second inertial measurement unit is affixed with respect to that second portion. A control circuit operably couples to those inertial measurement units and receives corresponding information regarding those portions of the tool. The control circuit can then process that received information to generate positional proprioception information as regards those monitored tool portions.
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 gas turbine engine includes a fan, a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining in part a working gas flowpath, the compressor section including a frame and defining a cavity forward of the frame, the compressor section further comprising a booster disposed in the working gas flowpath including a vane in fluid communication with the working gas flowpath and the cavity, the gas turbine engine defining a bypass passage over the turbomachine, and a heat exchanger disposed at least partially in the cavity, wherein the heat exchanger is in fluid communication with the booster in the working gas flowpath and is in fluid communication with the bypass passage.
An aerodynamic device defining a thickness direction is provided. The aerodynamic device configured to produce lift or thrust or configured to be a part of an aerodynamic system that produces lift or thrust. The aerodynamic device includes a cowl assembly that defines at least in part an airflow stream. The cowl assembly includes a first cowl and a second cowl moveable relative to the first cowl. The first cowl includes a plurality of first cowl indentations at an end of the first cowl. The second cowl defines an outer surface along the thickness direction and an inner surface along the radial direction. The second cowl includes a plurality of second cowl indentations complementary in shape to the plurality of first cowl indentations. The plurality of second cowl indentations are positioned locally on the outer surface of the second cowl or locally on the inner surface of the second cowl.
B64C 9/24 - Surfaces ou éléments de commande réglables, p. ex. gouvernes de direction formant des fentes à l'avant de l'aile par volet unique
B64D 29/06 - Fixation des nacelles, carénages ou capotages
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
A turbomachine engine including a high-pressure compressor, a high-pressure turbine, a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine, and a power turbine in flow communication with the high-pressure turbine. At least one of the high-pressure compressor, the high-pressure turbine, and the power turbine comprises a ceramic matrix composite (CMC) material. The turbomachine engine includes a low-pressure shaft coupled to the power turbine and characterized by a midshaft rating (MSR) between two hundred (ft/sec)1/2 and three hundred (ft/sec)1/2. The low-pressure shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine is configured to operate up to the redline speed without passing through a critical speed associated with a first-order bending mode of the low-pressure shaft.
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
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
Methods of forming a coating are presented. For example, a method for forming a coating on particles may include mixing an initial powder with a pressing medium, the initial powder comprising a plurality of core particles having an initial shell coating thereon; isostatic pressing the initial powder within the pressing medium to densify the initial shell coating on the plurality of core particles to form a pressed powder, the pressed powder comprising a densified shell coating on the plurality of core particles; and thereafter, removing the pressing medium from the pressed powder.
An analog system for edge artificial intelligence computing includes a first plurality of analog edge devices configured to receive an input analog signal and to output a first plurality of output analog signals, a second plurality of analog edge devices configured to receive the first plurality of output analog signals and to output a second plurality of output analog signal, and one or more memory devices in communication with the first plurality of analog edge devices and the second plurality of analog edge devices, and configured to store weight parameters, the weight parameters being adjustable based on time constants of the first plurality of analog edge devices or the second plurality of analog devices, or both. The second plurality of output analog signals are multiplied by the weight parameters to obtain a plurality of weighted analog signals.
Heat exchangers for gas turbine engines are described herein. An example heat exchanger includes a first plenum, a second plenum, and divider plates coupled to and extending between the first and second plenums. The divider plates are spaced apart from each other. Each of the divider plates has one or more internal passages for a fluid to flow between the first and second plenums. The heat exchanger also includes stiffeners between each adjacent pair of the divider plates, wherein at least one of the stiffeners is skewed relative to a connecting one of the divider plates.
A gas turbine engine includes a fan assembly, a turbo-engine encased within a turbo-engine cowl structure, a nacelle, and a thrust reverser system arranged, at least in part, in the nacelle. The turbo-engine includes a low-pressure compressor, a high-pressure compressor, an inter-compressor frame structure between the low-pressure compressor and the high-pressure compressor and including an outer frame portion having a cowl door engagement member on an upstream side of the outer frame portion, and a plurality of cowl doors defining the turbo-engine cowl structure. Each cowl door includes an inter-compressor frame engagement portion that engages with the cowl door engagement member of the inter-compressor frame structure, and a plurality of thrust reverser drag link connectors arranged on an outer side of the cowl door and arranged between the upstream side of the inter-compressor frame structure and a downstream side of the inter-compressor frame structure.
F02K 1/72 - Inversion du flux de la soufflante utilisant des volets inverseurs de poussée ou des portes montées sur le carter de la soufflante la partie arrière du carter de la soufflante étant mobile pour découvrir des ouvertures d'inversion de poussée dans le carter de la soufflante
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 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
F02C 9/52 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel avec la commande du flux du fluide de travail par prélèvement ou bipasse du fluide de travail
A gas turbine engine including a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section having a fuel nozzle including a fuel nozzle body defining an axis and having an inner surface defining a channel fluidly coupled to a combustion chamber, a support matrix located within the channel comprising a plurality of segments which intersect each other when viewed from aft, and a set of vortex generators located on the support matrix.
A system may include an image sensor. A system may include an actuator configured to cause a controlled movement of the image sensor relative to a target element, the controlled movement being based on an operating velocity of the target element relative to an initial position of the image sensor. A system may include a controller communicatively coupled to the image sensor, the controller configured to: identify the operating velocity, determine an activation time to activate the image sensor for a designated exposure time based on the operating velocity and the controlled movement; and activate the image sensor at the activation time.
H04N 23/695 - Commande de la direction de la caméra pour modifier le champ de vision, p. ex. par un panoramique, une inclinaison ou en fonction du suivi des objets
A gas turbine engine includes a turbomachine, a primary fan driven by the turbomachine, a secondary fan, a booster, and an outlet guide vane. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The secondary fan is located downstream of the primary fan within the inlet duct. The booster is located downstream of the secondary fan and includes a booster rotor blade and booster cowl that separates an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet. The outlet guide vane is positioned downstream of the secondary fan and upstream of the upper fan duct inlet or positioned within the upper fan duct.
A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, a booster located downstream of the secondary fan and comprising a booster rotor blade and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, and a flow blocker located at the lower fan duct inlet and movable from an open position to a closed position, wherein, in the closed position, the flow blocker blocks a flow through at least a portion of the lower fan duct inlet.
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
F02C 3/00 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail
A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, and a booster located downstream of the secondary fan and comprising a booster rotor blade, an inlet guide vane, and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet, the inlet guide vane located forward of the booster rotor blade.
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
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
