ПЕРЕПУСКНОЙ КАНАЛ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, СОДЕРЖАЩИЙ РЕШЕТКУ РПК С РАЗЛИЧНЫМИ УГЛАМИ УСТАНОВКИ

Внутренний корпус (2) промежуточного корпуса (1) двухконтурного газотурбинного двигателя. Перепускные лопатки (20) закреплены в канале (18) перепускного прохода на уровне выходного отверстия (6) наружной обечайки (5). При этом указанные перепускные лопатки (20) включают в себя от входа к выходу в направлении прохождения потока воздуха в пространстве (14) потока второго контура верхнюю по потоку лопатку (22), которая расположена рядом с верхней по потоку стенкой (18а) канала (18) перепускного прохода и которая содержит переднюю кромку (ВА), расположенную встречно потоку (F18) воздуха в канале (18) перепускного прохода, и среднюю линию (С) и нижнюю по потоку лопатку (24), которая расположена рядом с нижней по потоку стенкой (18b) канала (18) перепускного прохода и которая содержит переднюю кромку (ВА) и среднюю линию (С). При этом внутренний корпус (2) промежуточного корпуса отличается тем, что верхний по потоку острый угол () между плоскостью (Р) перепуска и касательной (Т) к средней линии ...

КАМЕРА СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, СОДЕРЖАЩАЯ ЗАХОДЯЩУЮ ДЕТАЛЬ С ОТВЕРСТИЕМ

Камера сгорания газотурбинного двигателя содержит по меньшей мере одну стенку, ограничивающую камеру сгорания и содержащую отверстие для прохождения заходящей детали. Указанная заходящая деталь содержит в своей части, находящейся внутри камеры сгорания, по меньшей мере одно отверстие, выполненное с возможностью создания воздушной пленки охлаждения зоны на выходе заходящей детали. Внутренняя поверхность указанной по меньшей мере одной стенки содержит две пластинки, расположенные параллельно друг другу и выполненные с возможностью создания воздушной пленки охлаждения указанной по меньшей мере одной стенки, при этом заходящая деталь находится между двумя пластинками. Изобретение направлено на увеличение срока службы камеры сгорания с максимальным уменьшением температурных градиентов и локальных горячих точек. 2 н. и 9 з.п. ф-лы, 7 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С ОТБОРОМ ПОТОКА СЖАТОГО ВОЗДУХА

Изобретение относится к газотурбинному двигателю, содержащему отбор потока сжатого воздуха, поступающего из компрессора. Газотурбинный двигатель, включающий в себя: выпускной коллектор (7), который содержит множество стоек (10), при этом пространство, разделяющее стойки, образует отверстия, в которых проходит воздушный поток внутреннего контура газотурбинного двигателя. Также газотурбинный двигатель включает в себя по меньшей мере один трубопровод, выполненный с возможностью отбора на одном из своих концов потока сжатого воздуха. При этом другой конец трубопровода соединен с по меньшей мере одним отверстием выпускного коллектора (7) таким образом, чтобы направлять отобранный воздушный поток в указанный воздушный поток внутреннего контура, причем во время своего захождения в отверстие указанный отобранный воздушный поток имеет число Маха, меньшее или равное 0,5. Изобретение позволяет снизить акустическое влияние, связанное с отбором потока и его повторным введением. Также изобретение позволяет ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ С ОТБОРОМ ПОТОКА СЖАТОГО ВОЗДУХА

КАНАЛ РАЗГРУЗОЧНОГО ПРОХОДА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, СОДЕРЖАЩИЙ РЕШЕТКУ РДК С РАЗЛИЧНЫМИ УГЛАМИ УСТАНОВКИ

Abgasanlage für eine Brennkraftmaschine

Abgasanlage (1) für eine Brennkraftmaschine, mit einem Abgasturbolader (2), dessen Turbinenrad mittels einer Bypassleitung (5) überbrückbar ist, wobei ein Durchströmungsquerschnitt einer der Bypassleitung (5) zugeordneten Ventilöffnung (7) mit einem runden oder einen runden Grundkörper (21) aufweisenden Ventilteller (8) einer Ventileinrichtung (6) einstellbar ist, dadurch gekennzeichnet, dass der Ventilteller (8) in seiner Geschlossenstellung derart angeordnet ist, dass eine Längsmittelachse (20) des runden Ventiltellers (8) oder des runden Grundkörpers (21) beabstandet zu der Längsmittelachse (10) der Ventilöffnung (7) vorliegt.

Abgasturbolader

Abgasturbolader (1), der einen Gaskanal (2) und einen Bypasskanal (3) aufweist, wobei die Verbindung von Gaskanal (2) und Bypasskanal (3) mit einem Schwenkklappenventil (5) verschließbar ist, dessen Ventilklappe (6) mit einem Ventilhebel (7) verbunden ist, der drehfest auf einer Welle (8) sitzt, wobei die Welle (8) in einem Gehäuse (4) des Abgasturboladers (1) gelagert ist und auf der Welle (8) drehfest ein Betätigungshebel (9) angeordnet ist, dessen von der Welle (8) abgewandter Endabschnitt aus dem Gehäuse (4) herausgeführt und außerhalb des Gehäuses (4) mit einer Betätigungseinrichtung (10) verbunden ist, dadurch gekennzeichnet, dass die Welle (8) beidseitig des Betätigungshebels (9) im Gehäuse (4) an einer ersten Lagerstelle (11) zwischen dem Ventilhebel (7) und dem Betätigungshebel (9) und an einer zweiten Lagerstelle (12) zwischen dem Betätigungshebel (9) und einem Außenende der Welle (8) gelagert ist.

Abgasturbine eines Abgasturboladers mit einer abgedichteten Wastegate-Ventileinrichtung sowie Abgasturbolader

Die Erfindung betrifft eine Abgasturbine (20) eines Abgasturboladers (100) mit einer Wastegate-Ventilanordnung (10), sowie einen Abgasturbolader (100) mit einer solchen Abgasturbine (20) .Die Abgasturbine (20) weist ein Turbinengehäuse (21) und eine dessen Gehäusewand (21a) durchdringende Lagerbohrung (23), mit einer darin fixiert angeordnet Lagerbuchse (1) auf, in der eine Ventilspindel (15) der Wastegate-Ventileinrichtung (10), vom Innenraum des Turbinengehäuses (21) durch die Lagerbohrung (23) der Gehäusewand (21a) nach außen, geführt und um ihre Spindelachse (16) drehbar gelagert ist. Dabei greift die Lagerbuchse (1) zumindest an einem ihren axialen Enden jeweils mit einem Stellhebel-Anschlussflansch (181) oder/und einem Ventilspindelabsatz (151), mittels zumindest einem, bezüglich der Spindelachse (16) umlaufenden, sich axial erstreckenden Dicht-Steg (6) und zumindest einer dazu komplementären Dicht-Nut (7), nach Art einer Labyrinth-Dichtung in axialer Richtung kämmend ineinander.

Fuel metering system

Louvre offtake arrangement

Diffuser for a turbocharger, with a geometry reducing stress due to thermal expansion

There is provided a diffuser for a turbine, comprising: a support configured to mount to a turbine housing; a diffuser body configured to receive fluid from an outlet of the turbine, the diffuser body defining a longitudinal axis and having a perimeter with a length measured in a plane normal to the longitudinal axis; and a bridge configured to connect the support to the diffuser body, wherein the connection between the bridge and the diffuser body is confined to a continuous portion of the perimeter of the diffuser body that is not more than around 50% of the total length of the perimeter of the diffuser body.

Gas turbine with external combustion, applying a rotating regenerating heat

Gas turbine with external combustion, applying a rotating regenerating heat exchanger.

Gas turbine with external combustion, applying a rotating regenerating heat exchanger.

GAS TURBINE ENGINE AND CONTROL METHOD

A method of controlling a gas turbine engine (10), the gas turbine engine (10) having in axial flow series a compressor (14), a combustor (16), a compressor-turbine (18) and an exhaust (30), the gas turbine capable of operating in at least a high output power range (65R), a medium-high output power range (82R), a medium output power range (67R), a medium-low output power range (70R, 70R') and a low output power range (72R). The method comprising the steps during the medium-high output power range (82R) varying the angle of the variable guide vanes (46) so that a third predetermined temperature (T3) of the combustor (16) is maintained, during the medium output power range (67R) the variable guide vanes (46) are closed and bleeding a gas from a downstream part (36) of the compressor (14) to an upstream part (38) of the compressor (14) so that a first predetermined temperature (T1) of the combustor (16) is maintained, during the medium-low output power range (70R, 70R') the variable guide ...

CHAMBRE DE COMBUSTION DE TURBOMACHINE COMPORTANT UNE PIECE PENETRANTE AVEC OUVERTURE

L'objet principal de l'invention est une chambre de combustion (1) de turbomachine, comportant au moins une paroi (3) délimitant la chambre de combustion (1) et comportant un orifice (30) de passage d'une pièce pénétrante (28), caractérisée en ce que la pièce pénétrante (28) comporte, dans sa partie située à l'intérieur de la chambre de combustion (1), au moins une ouverture (34) capable de créer un film d'air de refroidissement de la zone en aval de la pièce pénétrante (28).

JET ENGINE COLD AIR COOLING SYSTEM

Methods and devices for cooling systems (700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling system can include: a first precooler (210) recieving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a VGT cooling system turbine (240) downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) for bypassing the turbine can also be included.

MODULAR LOUVER SYSTEM

Louver systems for gas turbine bleed air systems are disclosed. An example louver system may include a bleed system discharge opening arranged to vent bleed air from a bleed flow conduit and a plurality of pivotable louvers disposed proximate the discharge opening, the pivotable louvers being pivotable between a shut position and an open position. In the shut position, individual louvers may at least partially obstruct the discharge opening. In the open position, individual louvers may at least partially control a direction of flow of the bleed air exiting the discharge opening.

GAS TURBINE AND METHOD FOR OPERATING A GAS TURBINE

Die Leistung einer erfindungsgemässen stationären Gasturbinenanlage bzw. deren Wirkungsgrad wird erhöht, indem der Luft-/Gasstrom – zwischen (Niederdruck)-Kompressoraustritt (1) und (Niederdruck)-Turbineneintritt (3) – in zwei oder mehrere parallele Teilströme aufgeteilt und der thermodynamische Prozess der Teilströme gegen höheren Druck und höhere Temperaturen hin angehoben wird. Dazu werden zwischen Niederdruck-Kompressor (1) und Niederdruck-Gasturbine (3), weitere Gasturbineneinheiten, jede bestehend aus Kompressor (1a), Verbrennungsraum (2a) und Gasturbine (3a), angeordnet, welche für den Betrieb auf hohem Druck- und Temperaturniveau ausgelegt sind. Die Achsen der Hochdruck-Gasturbineneinheiten können im Raum frei positioniert werden, müssen also nicht parallel zu oder identisch mit der Achse der Unterstufe sein, und können demzufolge optimal an die physischen Gegebenheiten der Maschine angepasst werden.

Druckdosenvorrichtung für Abgasturbolader sowie Verbrennungsmotor.

Die Erfindung betrifft eine Druckdosenvorrichtung zur pneumatisch-mechanischen Verstellung eines Waste-Gates oder der Turbinen-Schaufeln eines Abgasturboladers, umfassend ein Gehäuse mit einer darin angeordneten Membran, die mit einem Stellglied zur Übertragung einer mechanischen Stellkraft auf das Waste-Gate oder die Turbinen-Schaufeln verbunden ist, dadurch gekennzeichnet , dass im Bereich der Gehäusewand der Druckdose ein oder mehrere Kühlkanäle vorgehalten sind, die von einem externen Kühlmittelstrom zur dedizierten Kühlung der Druckdose durchströmbar sind.

Druckdosenvorrichtung für Abgasturbolader sowie Verbrennungsmotor.

Die Erfindung betrifft eine Druckdosenvorrichtung zur pneumatisch-mechanischen Verstellung eines Waste-Gates oder der Turbinen-Schaufeln eines Abgasturboladers, umfassend ein Gehäuse mit einer darin angeordneten Membran, die mit einem Stellglied zur Übertragung einer mechanischen Stellkraft auf das Waste-Gate oder die Turbinen-Schaufeln verbunden ist, dadurch gekennzeichnet , dass im Bereich der Gehäusewand der Druckdose ein oder mehrere Kühlkanäle vorgehalten sind, die von einem externen Kühlmittelstrom zur dedizierten Kühlung der Druckdose durchströmbar sind.

BASED ON TWO PINCH - POINTS CRITERION FOR OPTIMIZATION OF REGENERATIVE CYCLES RENKINA

BLEED VALVE ASSEMBLY FOR A GAS TURBINE ENGINE

Combined butterfly/spherical valve

By a turbine centrifugal pump fluid recirculation

HOUSING HUB FOR AIRCRAFT TURBOMACHINE COMPRISING AN AEROBIC COMPACT WITH INTEGRATED DEFLECTOR DOWNSTREAM FLANGE AT

TURBOMACHINE WITH TAKING A STREAM OF COMPRESSED AIR

GAS TURBINE ROTOR BLADE, AND GAS TURBINE

ACTUATOR SHAFT BOOT

Opening and closing device for high temperature gas passage

OPENING/CLOSING DEVICE FOR HIGH-TEMPERATURE GAS PASSAGE

In this exhaust damper (7), a high-temperature gas passage (8) and a valve (26) that opens and closes the high-temperature gas passage (8) are provided within a casing (18). The casing (18) of the exhaust damper (7) comprises a lining member (42) that is made from a metal sheet and that is exposed to the high-temperature gas passage (8), a cladding member (44) that is made from a metal sheet and that is exposed to the outside air, and a flexible insulating material (46) that is interposed therebetween. The lining member (42) engages with the insulating material (46) and is connected to the cladding material (44) via the insulating material (46) so as to be capable of movement relative to the cladding material (44).

VALVE DEVICE FOR A BLOW-OFF VALVE OF A TURBOCHARGER

The invention relates to a valve device (20) for a blow-off valve of a turbocharger (10), comprising a valve plate (22), by means of which a flow cross-section of a bypass channel (18) of the turbocharger (10) can be adjusted, wherein a gas is to bypass an impeller of the turbocharger (10) via said bypass channel. An end-face valve surface (32) is also formed by means of said valve plate, said valve surface (32) being formed in an asymmetrical manner.

TURBINE HOUSING, TURBOCHARGER, AND METHOD FOR MANUFACTURING TURBINE HOUSING

This turbine housing is formed from a light metal and accommodates a turbine wheel. This turbine housing includes: a bypass passage that extends so as to circumvent the turbine wheel, and enables communication between the upstream and downstream sections of the turbine wheel in the direction in which exhaust flows within the turbine housing; a wastegate valve that switches between opening and closing the bypass passage, and has a valve seat and a valve body capable of coming into contact with and separating from the valve seat; and a through hole that forms part of the bypass passage, and has an opening rim that forms the valve seat. The valve seat is reinforced by buildup welding.

RECIRCULATION OF FLUID THROUGH A TURBOMACHINE CENTRIFUGAL PUMP

A supply system for supplying a turbomachine with fluid. The supply system includes at least one centrifugal pump and a fluid recirculation branch. The fluid recirculation branch includes an inlet situated downstream of the centrifugal pump and an outlet situated upstream of the centrifugal pump or fluidically connected with a node situated upstream of the centrifugal pump, in such a way that at least one portion of the fluid circulates in the centrifugal pump. The fluid recirculation branch includes a valve situated between the inlet and the outlet, the valve including a shutter configured to open/close as a function of the temperature of the fluid.

Gas turbine engine

A gas turbine is provided for an aircraft comprising an engine core and a core flow path, a fan, a front drum cavity arranged radially inward of the core flow path, and a front bearing chamber. The front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor. The front drum inlet is through a seal, and the front drum outlet is through a spaced gap.

TURBINE ENGINE WITH A TURBO-COMPRESSOR

A turbine engine is provided that includes a turbo-compressor, a combustor section, a flow path and a recuperator. The turbo-compressor includes a compressor section and a turbine section. The combustor section includes a combustor and a plenum adjacent the combustor. The flow path extends through the compressor section, the combustor section and the turbine section. The recuperator is configured with the flowpath between the combustor section and the turbine section. An inlet duct to the recuperator is fluidly coupled with the flow path upstream of the plenum. An outlet duct from the recuperator is fluidly coupled with the plenum.

Gas turbine engine with active clearance control

A small gas turbine engine, such as is used to power a UAV, that includes at least one centrifugal compressor having an impeller with blades that form a gap between the blade tips and stationary shroud of the gas turbine engine, and where a resistance heating element is secured to or bonded to a compressor casing of the gas turbine engine in order to use heat to control the gap between the impeller blades and the stationary shroud. The resistance heating element is activated at cruise mode to move the shroud toward the impeller. Additionally or alternatively, the compressor casing is heated with bled-off compressed air to move the shroud toward the impeller. A capacitance tip clearance sensor can be mounted on the impeller shroud to monitor and control tip clearance in real time.

Gas-turbine engine with bleed-air tapping device

A gas-turbine engine with at least one compressor and at least one bleed-air tapping device, which includes an annular duct in a radially outer wall of a flow duct, and with an annular closing element, which is arranged in the region of the annular duct and can be moved in a substantially axial direction from a closed position to an open position, with the closing element having an annular flow divider projection which in the open position projects in the flow duct.

Passive stability bleed valve with adjustable reference pressure regulator and remote override capability

A bleed air valve comprises a housing that includes an inlet, an outlet, and a center portion to selectively provide a pneumatic flow path between the inlet and the outlet. A piston moves along a center guide in response to forces applied to the piston, where the piston is spring biased to an open position. The valve also comprises a pressure regulator valve that comprises a spring loaded poppet valve that includes a pressure regulator inlet that receives an inlet pressure which is regulated based upon force applied to a pressure regulator piston. A valve spring applies a force to the piston first surface to spring bias the bleed air valve to the open position. A solenoid valve receives a command signal and in response provides compressed air to a solenoid valve outlet that is in fluid communication with the pressure regulator inlet.

High Pressure Relief Valve Nozzle

A high pressure relief valve includes a nozzle that is received within an internal bore of a valve housing. The nozzle has a nozzle body with a bore defined by an outlet bore inner diameter that is less than an inlet bore inner diameter. The nozzle body has a nozzle face defined at the outlet end, and is defined between the outlet bore inner diameter and an outlet bore outer diameter spaced radially outwardly of the outlet bore inner diameter. The nozzle body includes a coupling feature held fixed relative to the nozzle body such that the coupling feature is not moveable relative to the nozzle body. The coupling feature is configured to couple the nozzle body to the closure sleeve that surrounds the outlet end. At least one retaining feature prevents the coupling feature from being uncoupled from the closure sleeve. 1. A high pressure relief valve comprising:a valve housing defining an internal bore and having a valve inlet configured to be in fluid communication with a pump outlet;a closure sleeve at least partially received within the internal bore, the closure sleeve comprising a sleeve body surrounding a center axis, wherein the sleeve body has an internal cavity that is enclosed at the downstream end and open at the upstream end;a piston received within the internal cavity, wherein the internal cavity is defined in part by a piston contact surface that is defined by an inner diameter, wherein the piston contact surface slides against an outer surface of the piston;a nozzle received within the internal bore of the valve housing, the nozzle comprising a nozzle body surrounding a center axis and defined by an overall length extending from an inlet end to an outlet end, wherein the nozzle body has a nozzle bore defined by an outlet bore inner diameter at the outlet end and an inlet bore inner diameter at the inlet end, the inlet bore inner diameter being greater than the outlet bore inner diameter, and wherein the nozzle body has a nozzle face defined at the outlet ...

Bypass valve assembly for turbine generators

A bypass valve assembly for use in turbine generators includes a valve body defining a central bore and a plurality of passageways. Each passageway has a smaller area at an inlet portion and a larger area at an outlet portion to define a flared passageway. A plurality of bypass valves is disposed within the plurality of passageways within the valve body. Each bypass valve includes a base portion and a nose portion, with each nose portion defining a predefined contoured surface area. At least a portion of the contoured surface area includes a wear coating disposed thereon. Optionally, the wear coating includes a plasma enhanced magnetron sputtering nanocoating.

Diffuser in wastegate turbine housings

A number of variations may include a turbine housing comprising a turbine body; an inlet passage and an outlet passage connected to the turbine body; a wastegate passage operatively connected to the outlet passage; a diffuser positioned within the outlet passage comprising at least one radial opening; wherein the first flow passage accepts fluid flow from the wastegate passage and the second flow passage accepts fluid flow from the turbine wheel; wherein a first end of the diffuser is attached to a first end of the turbine outlet and a second end of the diffuser is attached to a second end of the turbine outlet so that fluid flow from the first flow passage is directed into the second flow passage through the at least one radial opening before exiting the outlet passage, and wherein the at least one radial opening minimizes turbulence of fluid flow exiting the turbine housing.

Stator configuration for gas turbine engine

A stator configuration for a gas turbine engine including: a splitter segment, the splitter segment extending from a forward end to a rearward end; a forward most first stator extending radially outwardly from the splitter segment, the forward most first stator being completely located downstream from the forward end of the splitter segment; and a forward most second stator extending radially inwardly from the splitter segment, the splitter segment, the forward most first stator and the forward most second stator being formed as a single, integrally formed structure, the forward most first stator positioned closer to the forward end relative to a distance between the forward most second stator and the forward end and the forward most second stator positioned closer to the rearward end relative to a distance between the forward most first stator and the rearward end.

DUCT STRUCTURE FOR AIRCRAFT PROPULSION SYSTEM

An apparatus is provided for an aircraft propulsion system. This apparatus includes a duct structure, and the duct structure includes a transition duct, an inlet duct and a bypass duct. The duct structure is configured as a monolithic body. The transition duct includes an inlet, a first outlet and a second outlet. The transition duct extends longitudinally along a longitudinal centerline from the inlet to the second outlet. The first outlet is arranged longitudinally along the longitudinal centerline between the inlet and the second outlet. A centerline axis of the first outlet is angularly offset from the longitudinal centerline. The inlet duct extends longitudinally along the longitudinal centerline to the inlet. The bypass duct extends longitudinally along the longitudinal centerline from the second outlet.

WASTE GATE VALVE

In a wastegate valve which is provided in the bypass path bypassing the turbine of the turbocharger in the exhaust gas path and which opens and closes the bypass path, the wastegate valve is provided with a valve seat which is formed in a plane perpendicular to or tilted with an inclination angle with respect to an axial direction of the bypass path, and a valving element which is pivotable around a pivot point which has a relationship of 0°<β<90° with respect to the plane including the valve seat where β is an inclination angle, the valving element being moved away from or closer to the valve seat by pivotation of the valving element to open or close the valve.

Gas turbine engine

Diffuser for a turbocharger, with a geometry reducing stress due to thermal expansion

Blade for a gas turbine engine

Turbocharger

A turbocharger comprises a compressor and a volume 29 behind the compressor impeller 7, which is in selective fluid communication with a location 37 downstream of the compressor via a connecting channel 47. A moveable element 53 is operable to block the channel 47 when the pressure at the downstream location 37 is sufficiently large. This arrangement ensures an adequate pressure in the volume 29 to prevent lubricant leakage, whilst maintaining efficiency by only bleeding compressed air when needed because the pressure in the volume is low, e.g. during low speed and low pressure operation.

A fuel staging system

PASSIVE BLADE TIP CLEARANCE CONTROL SYSTEM FOR GAS TURBINE ENGINE

The present disclosure relates to a gas turbine engine including a turbine wheel mounted for rotation about a central axis and a turbine shroud ring mounted radially outward from the turbine wheel. The turbine wheel includes a plurality of blades that are spaced apart radially from the turbine shroud ring to establish a blade tip clearance gap. The gas turbine engine further includes a blade tip clearance control system that passively controls the size of the clearance gap based on engine operation.

GAS TURBINE ENGINE FLUID COOLING SYSTEMS AND METHODS OF ASSEMBLING THE SAME

A fluid cooling system for use in a gas turbine engine including a fan casing circumscribing a core gas turbine engine includes a heat source configured to transfer heat to a heat transfer fluid and a primary heat exchanger coupled in flow communication with the heat source. The primary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the fan casing. The fluid cooling system also includes a secondary heat exchanger coupled in flow communication with the primary heat exchanger. The secondary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the core gas turbine engine. The fluid cooling system also includes a bypass mechanism coupled in flow communication with the secondary heat exchanger. The bypass mechanism is selectively moveable based on a temperature of a fluid medium to control a cooling airflow through the secondary heat exchanger.

AUXILIARY POWER UNIT BLEED CLEANING FUNCTION

An auxiliary power unit is operable to provide bleed air to a vehicle system. A method includes diverting substantially all of the bleed air to an exhaust for a selected time period commencing with startup of the auxiliary power unit. After the selected time period, at least a portion of the bleed air is diverted to the vehicle system.

METHOD AND SYSTEM FOR MITIGATION OF CAVITY RESONANCE

A turbofan engine includes a core engine, a fan, a fan bypass duct partially surrounding the core engine and the fan, and a bleed system. The bleed system includes a first bleed circuit configured to bleed pressurized air from the core engine and channel the flow to a first circuit of a heat exchanger, and a second bleed circuit configured to bleed fan air from the fan bypass duct and channel the flow to a second circuit of the heat exchanger. The second bleed circuit includes a bleed duct including a duct inlet and a duct outlet coupled in flow communication with the heat exchanger through a valve. The bleed duct also includes an acoustic suppression conduit extending from the bleed duct upstream of the valve to the fan bypass duct and sized to suppress pressure oscillations inside the second bleed circuit when the valve is at least partially closed.

GAS TURBINE EXTERNAL COMBUSTION WITH ROTATING REGENERATIVE HEAT EXCHANGER

SYSTEM AND METHOD OF STARTING PLANT GENERATION OF POWER

With the compressed air flow of collection of turbine

AN AIR INTAKE FOR A TURBOPROP

TURBOMACHINE COMBUSTION CHAMBER COMPRISING A WORKPIECE PENETRATING WITH OPENING

TURBOJET LOW NOISE FAN

HANDLE AIR INLET FOR A MOTOR

Manche d'entrée d'air (32) pour un turbopropulseur d'aéronef, comprenant un canal (54) de prélèvement d'air orienté sensiblement selon un premier axe (A) et un canal (56) d'alimentation en air d'un compresseur, qui est orienté sensiblement selon un second axe (B) à distance du premier axe et sensiblement parallèle au premier axe, ledit canal de prélèvement et ledit canal d'alimentation étant raccordés ensemble par un canal intermédiaire (60) de forme générale en S, vu de côté, caractérisée en ce que ledit canal intermédiaire comprend sur au moins une de ses parois des moyens d'aspiration d'air dans la veine du canal intermédiaire, lesdits moyens d'aspiration (64, 66) étant situés et/ou configurés pour aspirer une couche limite de ladite veine.

A FUEL SYSTEM OF A TURBOMACHINE

Circuit de carburant d'une turbomachine, ce circuit comprenant : - une vanne de retour carburant (FRV) reliée au circuit principal (102) de carburant et à un réservoir (110), la vanne (FRV) pouvant adopter une première et une deuxième position ouverte, distinctes, et une position fermée, - deux lignes hydrauliques primaires (120, 130) reliant la vanne (FRV) au circuit principal (102) et comprenant, respectivement, des premier et deuxième filtres (125, 135) à travers lesquels passe le carburant lorsque la vanne (FRV) est dans sa première position ouverte, - deux lignes hydrauliques secondaires (140, 150) qui relient la vanne (FRV) au circuit principal (102) et qui sont positionnées par rapport aux premiers filtres (125, 135) de telle sorte que la circulation de carburant dans ces lignes secondaires (140, 150) contribue, respectivement, au nettoyage des premier et deuxième filtres (125, 135), le carburant circulant dans les lignes secondaires (140, 150) lorsque la vanne (FRV) est dans sa ...

INTERMEDIATE CASING FOR A JET ENGINE

VALVE FOR THE CONTROL OF THERMAL EXCHANGES OF AN OIL CIRCUIT

La vanne (2) pour la régulation en température d'un débit d'huile (BP1), ladite vanne (2) comprenant une première voie (25) et une seconde voie (27) d'entrées et une voie de sortie (8), la seconde voie (27) d'entrée et la voie de sortie (8) étant aptes à coopérer avec un régulateur (E2) de température. L'une des voies (25, 27) d'entrées comporte un régulateur (50) du débit d'huile (BP1), dont le pilotage est assuré au moyen d'un calculateur par la génération d'une consigne électrique respectant une loi de pilotage configurée au sein du calculateur, ladite loi de pilotage étant asservie par un capteur en température du débit d'huile (BP1), la consigne électrique commandant l'ouverture ou la fermeture du régulateur (50), ladite loi de régulation générant une alternance d'états d'ouverture et de fermeture du régulateur (50) de débit de manière à assurer sur une période donnée une température moyenne souhaitée du débit d'huile.

Single piece part i.e. casting part, for intermediate casing hub of e.g. turbojet engine, of aircraft, has deflecting surface whose radial internal end partially defines separation nozzle, where surface is extended to external end

L'invention concerne une pièce (70) réalisée d'un seul tenant pour moyeu (36) de carter intermédiaire de turbomachine d'aéronef, comprenant : - au moins une partie d'un flasque aval (42) du moyeu ; - au moins une partie d'une virole interne (38) du moyeu comprenant au moins un bec de séparation (60) en partie défini par une surface de délimitation externe (62) d'un flux primaire (22) de turbomachine. Selon l'invention, la pièce comprend de plus une surface déflectrice (65) pour le guidage de l'air, cette surface comprenant une extrémité radiale interne (65a) définissant en partie le bec (60), et se prolongeant jusqu'à une extrémité radiale externe (65b) formée par une surface amont du flasque aval (42).

액추에이터 샤프트 부트

... 웨이스트 게이트로 된 터보차저는 모두 배기가스 에너지의 터빈휠 우회를 제어하는 웨이스트 게이트 밸브를 조작하는 액추에이터들을 사용한다. 상기 액추에이터들 내의 상기 다이어프램들은 외부 물질에 의한 손상에 영향을 받기 쉽다. 이러한 손상은 상기 액추에이터 샤프트 주위에 부트를 추가하여, 액추에이터의 수명에 해로울 수 있는 외부 물질 및 유체들의 유입을 방지함으로써 최소화된다.

GAS TURBINE

louver system, gas turbines and method to operate a flush system

SYSTEM AND METHOD FOR ENGINE OPERATION IN A MULTI-ENGINE AIRCRAFT

Methods and systems for operating a gas turbine engine in a multi-engine aircraft are described. The method comprises operating the gas turbine engine in a standby mode to provide substantially no motive power to the aircraft when another engine of the multi-engine aircraft is operated in an active mode to provide motive power to the aircraft, transitioning the gas turbine engine from the standby mode to the non-standby mode, and applying pulse width modulation to an air switching system of the gas turbine engine while transitioning the gas turbine engine from the standby mode to the non-standby mode. 1. A method for operating a gas turbine engine in a multi-engine aircraft , the method comprising:operating the gas turbine engine in a standby mode to provide substantially no motive power to the aircraft when another engine of the multi-engine aircraft is operated in an active mode to provide motive power to the aircraft;transitioning the gas turbine engine from the standby mode to the non-standby mode; andapplying pulse width modulation to an air switching system of the gas turbine engine while transitioning the gas turbine engine from the standby mode to the non-standby mode.2. The method of claim 1 , further comprising monitoring at least one of pressure and temperature while applying pulse width modulation to the air switching system.3. The method of claim 2 , wherein applying the pulse width modulation comprises adapting the pulse width modulation in real-time based on the at least one of pressure and temperature as monitored.4. The method of claim 3 , wherein adapting the pulse width modulation comprises determining a duration of the pulse width modulation.5. The method of claim 3 , wherein adapting the pulse width modulation comprises determining a duration of time the air switching system stays in a first position and a second position for each cycle of a pulse width modulation signal.6. The method of claim 1 , wherein applying pulse width modulation ...

INTEGRATED ENVIRONMENTAL CONTROL AND BUFFER AIR SYSTEM

An environmental control system for an aircraft includes a higher pressure tap associated with a higher compression location in a main compressor section. The higher pressure tap leads into a turbine section of a turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive a compressor section of the turbocompressor. A combined outlet receives airflow from a turbine outlet and a compressor outlet intermixing airflow and passing the mixed airflow downstream to be delivered to an aircraft system. A buffer air outlet communicates airflow to an engine buffer air system.

Steam turbine control device

A steam turbine control device has first and second valves, first and second valve controllers, and a valve control adjuster. The first valve is provided in a first steam supply path connected to a steam turbine. The second valve is provided in a second steam supply path connected to a lower-pressure side of the steam turbine while bypassing the first valve from on the first steam supply path. The first valve controller controls an opening degree of the first valve based on flow rate information A designating a flow rate of steam to be sent to the steam turbine. The second valve controller controls an opening degree of the second valve based on the flow rate information A. The valve control adjuster adds adjustment to control of the opening degree of the second valve by the second valve controller.

Turbine section of high bypass turbofan

A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio ratio is below about 170.

Mixing bleed and ram air using an air cycle machine with two turbines

An air cycle machine for an environmental control system for an aircraft is provided. The air cycle machine includes a compressor configured to compress a first medium, a turbine configured to receive second medium, a mixing point downstream of the compressor and downstream of the turbine; and a shaft mechanically coupling the compressor and the turbine.

Generator temperature management for throttle loss recovery systems

Turbine assemblies, loss recovery systems, and related fabrication methods are provided for managing temperatures associated with an electrical generator. One exemplary turbine assembly suitable for use in a loss recovery system includes a wheel configured to rotate in response to a portion of a fluid flow bypassing a flow control valve, a generator including a stator assembly disposed about a rotor coupled to the wheel to rotate in response to rotation of the wheel, a conductive structure in contact with the stator assembly, and an insulating structure radially encompassing the conductive structure and the generator. The conductive structure accesses at least a portion of the fluid flow bypassing the flow control valve and impacting the wheel, thereby providing thermal coupling between the stator assembly and the bypass fluid flow to transfer heat from the stator assembly to the bypass fluid flow via the conductive structure.

AIRCRAFT TURBOMACHINE HAVING AN AIR INLET OF VARIABLE SECTION

A turbomachine for an aircraft comprising a fan duct delimited by a wall and through which an air stream flows from upstream to downstream, and an air passage arranged in the wall comprising an air inlet opening flush with the wall, the air passage being designed to draw part of the air flow of the fan duct through the air inlet opening. To control the amount of air entering the air passage, a flap is rotatably mounted on the wall of the fan duct, about an axis of rotation disposed downstream of the air inlet opening, between an open position, in which the flap partially closes the air inlet opening and leaves free the fan duct downstream of the air inlet opening, and a closed position, in which the flap leaves open the air inlet opening and partially closes the fan duct downstream of the air inlet opening. 1. A turbomachine for an aircraft , comprising:a fan duct delimited by a wall and through which a stream of air circulates from upstream to downstream,an air passage arranged in the wall and comprising an air inlet opening flush with the wall, the air passage being arranged to draw part of the flow of the air of the fan duct through said air inlet opening, anda flap mounted rotatably on the wall of the fan duct, about an axis of rotation disposed downstream of the air inlet opening, between an open position, in which the flap partially closes the air inlet opening and leaves free the fan duct downstream of said air inlet opening, and a closed position, in which the flap leaves open the air inlet opening and partially closes the fan duct downstream of said air inlet opening.2. The turbomachine as claimed in claim 1 , wherein a face of the flap oriented toward the fan duct is flush with the wall of said fan duct when the flap is in the closed position.3. The turbomachine as claimed in claim 1 , further comprising blocking means when arranged in a first locking position claim 1 , said means lock the flap in the closed position claim 1 , and when arranged in a second ...

Diffuser for a radial compressor

The present disclosure relates to a diffuser for a radial compressor. The diffuser may comprise a diffusor duct portion formed by first and second side walls that are arranged so as to diverge at least partially from one another in a direction of flow, a blade ring having a number of blades arranged at least partially in the diffusor duct portion with each blade having a pressure side and a suction side delimited by a blade leading edge and by a blade trailing edge of the respective blade, a number of pressure equalizing openings incorporated into at least one of the first and second side walls of the diffuser duct portion in a region where the first and second side walls diverge from one another with each of the pressure equalizing openings being arranged between the pressure side of one blade and the suction side of an adjacent blade of the blade ring, and a first annular duct arranged behind the pressure equalizing openings and fluidically connected to the diffuser duct portion via at ...

GAS TURBINE ENGINE AND METHOD OF ASSEMBLING THE SAME

A gas turbine engine having a forward thrust mode and a reverse thrust mode is provided. The gas turbine engine includes a variable pitch fan configured for generating forward thrust in the forward thrust mode of the engine and reverse thrust in the reverse thrust mode of the engine. The engine also includes a fan cowl surrounding the variable pitch fan, wherein the fan cowl forms a bypass duct for airflow generated by the fan. The fan cowl includes an aft edge that defines a physical flow area of the bypass duct, and a deflection device configured for deflecting airflow near the aft edge, wherein the deflection device is configured for operation in the reverse thrust mode of the engine. The physical flow area of the bypass duct at the aft edge remains the same in the forward thrust mode of the engine and in the reverse thrust mode of the engine.

Steam turbine, partition member, and method for operating steam turbine

A steam turbine includes: a partition section that partitions a high-pressure stage and a low-pressure stage; and a pressure regulation valve that regulates a pressure of extraction steam. The pressure regulation valve includes: a plurality of flow rate regulation valves; and a plurality of flow path compartments that correspond to the respective flow rate regulation valves and that communicate with the low-pressure stage side relative to the partition section through respective nozzle holes. The plurality of flow path compartments are arranged over the entire partition section in a circumferential direction in a region including an outer peripheral side of the pressure regulation valve relative to the partition section as a whole. The partition section includes a bypass passage that makes the high-pressure stage side and the low-pressure stage side communicate with each other without passing through the pressure regulation valve.

Thermal management system

A method for thermal management for an aircraft includes extracting a flow of compressed fluid from a compressor section of a propulsion system. The flow of compressed fluid is passed through an anti-ice system. The flow of compressed fluid flows from the anti-ice system to a turbine. The flow of compressed fluid is expanded across the turbine. The expanded flow of compressed fluid then flows to thermal communication with a thermal load.

Gas turbine engine compressor particulate offtake

In general, techniques are described by which to provide a turbine engine compressor particulate offtake. A gas turbine engine comprising a compressor may perform the techniques. The compressor may comprise a plurality of compressor stages, where each compressor stage comprises a circumferential row of stator vanes and a rotor. The compressor may also include an outer circumferential casing that defines a fluidic opening within or between adjacent compressor stages. The fluidic opening may be configured to receive particulate flowing and air through the compressor and output the particulate and air outside of the gas turbine engine. The gas turbine engine also includes a combustor in series flow downstream of the compressor, and a turbine in series flow downstream of the combustor.

ACOUSTICALLY OPTIMIZED DISCHARGE LINE GRID WITH CHANNELS

AIRCRAFT TURBINE ENGINE ASSEMBLY COMPRISING A PASSIVE VALVE FOR BY-PASSING A FUEL/OIL HEAT EXCHANGER

КАМЕРА СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, СОДЕРЖАЩАЯ ЗАХОДЯЩУЮ ДЕТАЛЬ С ОТВЕРСТИЕМ

СИСТЕМА УПРАВЛЕНИЯ ПОТОКОМ С СЕТЬЮ ПАРАЛЛЕЛЬНО СОЕДИНЕННЫХ ТОПЛИВНЫХ КАНАЛОВ

Aufgeladene Brennkraftmaschine mit in Reihe angeordneten Abgasturboladern

Die Erfindung betrifft eine aufgeladene Brennkraftmaschine mit einem Ansaugsystem zum Zuführen von Ladeluft, einem Abgasabführsystem (1) zum Abführen des Abgases, mindestens einer Abgasrückführung, und mindestens zwei in Reihe geschalteten Abgasturboladern (2, 4), die jeweils eine im Abgasabführsystem (1) angeordnete Turbine (2a, 4a) und einen im Ansaugsystem angeordneten Verdichter (3, 5) umfassen und von denen ein erster Abgasturbolader (2) als Niederdruckstufe (2) und ein zweiter Abgasturbolader (4) als Hochdruckstufe (4) dient, wobei die zweite Turbine (4a) des zweiten Abgasturboladers (4) stromaufwärts der ersten Turbine (2a) des ersten Abgasturboladers (2) angeordnet ist. Es soll eine aufgeladene Brennkraftmaschine der oben genannten Art bereitgestellt werden, mit der die aus dem Stand der Technik bekannten Nachteile überwunden werden und deren Aufladeverhalten verbessert ist. Erreicht wird dies durch eine Brennkraftmaschine der oben genannten Art, die dadurch gekennzeichnet ist, ...

Zentrifugalverdichter

Ein Zentrifugalverdichter weist einen Motor, der eine Drehwelle eines Verdichterlaufrads dreht, ein Motorgehäuse, das den Motor aufnimmt, ein Verdichtergehäuse, das das Verdichterlaufrad aufnimmt und eine Ansaugöffnung und eine Abgabeöffnung aufweist, eine Luftauslassöffnung, die in einer Strömungsrichtung in dem Verdichtergehäuse näher an der Abgabeöffnung vorgesehen ist als das Verdichterlaufrad, eine Kühlgasleitung, welche mit der Luftauslassöffnung verbunden ist und durch welche ein Teil eines verdichteten Gases tritt, das durch das Verdichterlaufrad verdichtet wird, eine Kühlflüssigkeitsleitung, von welcher zumindest ein Teil in dem Motorgehäuse vorgesehen ist und durch welche eine Kühlflüssigkeit tritt, deren Temperatur niedriger ist als die Temperatur des verdichteten Gases, und einen Wärmetauscher auf, der an der Kühlgasleitung und der Kühlflüssigkeitsleitung angeordnet ist.

Abgasturbolader

Die Erfindung betrifft einen Abgasturbolader (1), welcher eine Turbinengehäuse (3) ein Verdichtergehäuse (5) und ein dazwischen angeordnetes Lagergehäuse (10) aufweist, wobei das Turbinengehäuse (3) und das Lagergehäuse (10) über einen Turbinengehäuseflansch (3a) und einen Lagergehäuseflansch (10a), die sich in einem radialen Randbereich in axialer Richtung (A) überlappen, miteinander verbunden sind, wobei zur axialen Festlegung des Turbinengehäuses (3) relativ zum Lagergehäuse (10), unter gleichzeitiger Vorspannung einer dazwischen angeordneten Federvorrichtung (11), in jeweils zueinander korrespondierende Aussparungen (12, 13) des Turbinengehäuseflansches (3a) und des Lagergehäuseflansches (10a) Verbindungselemente (14) im Wesentlichen senkrecht zur axialen Richtung (A) so eingesetzt sind, dass diese sich vom Turbinengehäuseflansch (3a) bis in den Lagergehäuseflansch (10a) erstrecken.

WASTEGATEANORDNUNG

Eine Wastegateanordnung zur Steuerung des Abgasstroms umfasst ein Ventilelement mit einem Ventilkörper und einem Ventilschaft. Die Wastegateanordnung umfasst ferner eine Spindel mit einem Kopf, der eine Öffnung definiert und eine flache Fläche umfasst. Die Wastegateanordnung umfasst ferner eine Scheibe, die mit dem Ventilschaft gekoppelt und von der Spindel beabstandet ist, zur Sicherung der Spindel an dem Ventilschaft. Die Scheibe definiert eine untere Scheibenfläche, die zu der flachen Fläche des Spindelkopfs weist, wobei die untere Scheibenfläche einen flachen Bereich und einen angefasten Bereich umfasst. Ein Vorspannglied ist zwischen der flachen Fläche der Spindel und dem flachen und dem angefasten Bereich der Scheibe angeordnet. Der angefaste Bereich erstreckt sich schräg zu dem flachen Bereich zur Minimierung des Kontakts zwischen der Scheibe und dem Vorspannglied über den flachen Bereich hinaus.

A valve assembly

Secondary flow control

Oil cooling system

Turbine arrangement

A turbine bypass valve comprising a valve member 241, operable to open or close an aperture 23 between the upstream and downstream sides of the turbine, and moved by a first actuator arm 25. A second actuator including a piston 31 connected to the gas upstream of the valve is forced by the pressure of the gas against the valve member, thereby urging the valve towards the closed position. The valve member is thereby balanced and may be closed with less force by the first actuator arm 25. The piston may be larger, smaller or the same size as the aperture such that the valve member is urged towards the closed position, the open position or there is no net force on the valve member. The piston may have a curved lower portion 312 abutting the valve and may be a sphere. The turbine may be a turbocharger for an internal combustion engine.

POPPET VALVE SYSTEM AND PROCESS

Aspects herein include a valve to provide regulated fluid flow. The valve comprises a valve housing having an inlet and an outlet. The valve further comprises a valve seat disposed between the inlet and the outlet of the valve. The valve seat has a seat opening defined by a seat opening dimension and is fixed in relation to the valve housing. The valve further comprises a poppet disposed between the valve seat and the outlet of the valve, the poppet having a seat face opposing the valve seat. The seat face tapers from a poppet large dimension larger than the seat-opening dimension disposed toward the inlet end to a poppet small dimension smaller than the seat-opening dimension disposed toward the outlet end. The valve further comprises a plunger operatively coupled with the poppet as well as a solenoid within the valve chamber operatively coupled with the plunger.

MULTI-STAGE PUMP WITH ENHANCED THRUST BALANCING FEATURES

A multi-stage pump featuring first and second stages, each stage having an impeller arranged on a rotor of the pump, each impeller having a hub-side and an eye-side, and each impeller configured to pump a liquid through the pump that applies an axial thrust load caused by a pressure difference in an axial direction from the hub-side to the eye-side of each impeller; and a first and second stage pump casing, each casing configured to form a casing enclosure to contain components of the first stage and the second stage, including each impeller, and configured with one or more pump casing openings formed therein and passing thru the pump casing to leak at least some liquid being pumped from inside to outside the casing enclosure to reduce substantially the axial thrust load caused by the pressure difference in the axial direction from the hub-side to the eye-side of each impeller.

TURBOMACHINE WITH COAXIAL PROPELLERS

The invention relates to a turbomachine with a longitudinal axis, comprising two, respectively upstream (122) and downstream, coaxial outer propellers (122), characterised in that at least some of the blades (148) of the upstream propeller (122) comprise at least one internal air circulation chimney (150) that communicates with air-bleeding openings (152) in the boundary layers of the blades (148), and communicates with air outflow openings (158) on the radially outer end thereof, the air-bleeding openings (152) leading to opening inlets (152a) on the passive surfaces (156) of the blades (148), the inlets (152a) of the air-bleeding openings being radially arranged in an area (H1) contained between 10% and 45% of the radial dimension (H2) of the blades (148), measured above and from the radial height of the blades for which the tangent of the leading edge (138) of the blades is orthogonal to the longitudinal axis, and the inlets (152a) of the air bleeding openings being arranged in an area ...

BYPASS TURBOMACHINE

L'invention concerne un ensemble pour turbomachine d'axe longitudinal comprenant une première paroi (24) annulaire, des panneaux (38) étant agencés autour de l'axe longitudinal (A) et s'étendant en vis-à-vis radial de ladite première paroi annulaire (24) de manière à former une surface d'écoulement d'un flux d'air, chaque panneau (38) étant solidarisé à la première paroi (24) annulaire par au moins un organe de fixation (72) traversant un orifice du panneau (38) et solidarisé à la première paroi (24) annulaire au moyen d'une douille et d'un plot formant une entretoise.

Wastegate valve

In a wastegate valve which is provided in the bypass path bypassing the turbine of the turbocharger in the exhaust gas path and which opens and closes the bypass path, the wastegate valve is provided with a valve seat which is formed in a plane perpendicular to or tilted with an inclination angle with respect to an axial direction of the bypass path, and a valving element which is pivotable around a pivot point which has a relationship of 0°<β<90° with respect to the plane including the valve seat where β is an inclination angle, the valving element being moved away from or closer to the valve seat by pivotation of the valving element to open or close the valve.

Spherical-link end damper system with near constant engagement

A link includes a link body with two ends, a ring bore with a ring bore axis and a bearing, a mount bore with a mount bore axis and a bearing. The link also has an end curvature at the end having the ring bore wherein the curvature axis is substantially perpendicular to the ring bore axis.

Turbine and turbocharger

A turbine includes: a turbine impeller accommodated in an accommodation space; two turbine scroll flow paths connected to the accommodation space; a first wastegate flow path (wastegate flow path) opened to one of the turbine scroll flow paths (second turbine scroll flow path) and separated from the other turbine scroll flow path (first turbine scroll flow path), and a valve for opening and closing the first wastegate flow path.

Mixing bleed and ram air using an air cycle machine with two turbines

An air cycle machine for an environmental control system for an aircraft is provided. The air cycle machine includes a compressor configured to compress a first medium, a turbine configured to receive second medium, a mixing point downstream of the compressor and downstream of the turbine; and a shaft mechanically coupling the compressor and the turbine.

MIXING BLEED AND RAM AIR USING AN AIR CYCLE MACHINE WITH TWO TURBINES

An air cycle machine for an environmental control system for an aircraft is provided. The air cycle machine includes a compressor configured to compress a first medium, a turbine configured to receive second medium, a mixing point downstream of the compressor and downstream of the turbine; and a shaft mechanically coupling the compressor and the turbine. 1. An air cycle machine for an environmental control system for an aircraft , the air cycle machine comprising:a compressor configured to compress a first medium;a turbine configured to receive second medium;a mixing point downstream of the compressor and downstream of the turbine;a shaft mechanically coupling the compressor and the turbine;a second turbine mounted on the shaft and configured to expand the first medium; anda fan driven by a motor.2. The air cycle machine of claim 1 , further comprising the fan on a second shaft.3. The air cycle machine of claim 2 , wherein the fan is located at a first end of the second shaft.4. The air cycle machine of claim 3 , wherein a third turbine is located at a first end of the shaft.5. The air cycle machine of claim 4 , further comprising the fan located at a second end of the shaft.6. The air cycle machine of claim 3 , wherein the second turbine is configured to receive a third medium claim 3 , andwherein the third medium is cabin discharge air.7. The air cycle machine of claim 1 , wherein the first medium comprises fresh air claim 1 , andwherein the second medium comprises bleed air.8. An air conditioning system for an aircraft comprising:a compressor configured to compress a first medium;a turbine configured to receive a second medium;a mixing point downstream of the compressor and downstream of the turbine; anda shaft mechanically coupling the compressor and the turbine;a second turbine mounted on the shaft and configured to expand the first medium; anda fan driven by a motor.9. The air conditioning system of claim 8 , further comprising the fan on a second shaft.10. ...

Activation Control Apparatus

An object of the invention is to provide an activation control apparatus by which a steam turbine can be activated safely at a high speed in response to a state of a power generation plant. In an activation control apparatus for a power generation plant including a heat source apparatus that heats low-temperature fluid by a heat source medium to generate high-temperature fluid; steam generation equipment that generates steam by thermal exchange with the high-temperature fluid; a steam turbine that is driven by the steam; and an adjustment apparatus that adjusts a plant operation amount, the activation control apparatus comprises: a thermal effect amount prediction calculation device that calculates at least a prediction value for a thermal effect amount for use for activation control of the steam turbine; a changeover device that decides, based on a state value of the power generation plant, the sensitivity of the thermal effect amount to a variation of the plant operation amount and outputs a changeover signal for a control mode for the thermal effect amount in accordance with the sensitivity; and an adjustment device that calculates, based on the changeover signal, the plant operation amount so as not to exceed a predetermined limit value. 1. An activation control apparatus for a power generation plant including:a heat source apparatus configured to heat low-temperature fluid by a heat source medium to generate high-temperature fluid;steam generation equipment configured to generate steam by thermal exchange with the high-temperature fluid;a steam turbine configured to be driven by the steam; andan adjustment apparatus configured to adjust a plant operation amount; the activation control apparatus comprising:a thermal effect amount prediction calculation device configured to calculate a prediction value for at least one thermal effect amount to be used for activation control of the steam turbine;a changeover device configured to decide, based on a state value of ...

DEVICES AND METHODS FOR GUIDING BLEED AIR IN A TURBOFAN ENGINE

Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss. 1. A device for guiding bleed air into a bypass duct of a turbofan engine having a central axis , the device comprising:a body defining a flow-guiding surface having opposite first and second ends defining a span of the flow-guiding surface around the central axis, the flow-guiding surface extending between a radially-inner edge of the body and a radially-outer edge of the body relative to the central axis; anda side wall adjacent the first end of the flow-guiding surface of the body, the side wall extending at least partially axially relative to the central axis, the side wall extending from a first position radially inwardly of the radially-inner edge of the body to a second position radially outwardly of the radially-inner edge of the body relative to the central axis.2. The device as defined in claim 1 , wherein the second position is adjacent the radially-outer edge of the body.3. The device as defined in claim 1 , wherein the side wall is substantially planar.4. The device as defined in claim 3 , wherein the side wall is non-parallel to a radial direction relative to the central axis.5. The device as defined in claim 1 , wherein the side wall is curved.6. The device as defined in claim 1 , wherein the side wall has a Bellmouth profile when viewed along the central axis.7. The device as defined in claim 1 , wherein the side wall has a unitary construction with the body.8. The device as defined in claim 1 , comprising a baffle disposed axially of the body to define a bleed air passage between the baffle and the flow-guiding surface of the body claim 1 , wherein a gap is defined between the side wall and the baffle.9. The device as defined in ...

INTERMEDIATE CASING FOR A TURBOFAN ENGINE

An intermediate casing comprising an inner annular hub, an outer annular barrel and an annular part for separating flows situated between the hub and the outer barrel. A primary stream is delimited between the hub and the separation part. A secondary stream is delimited between the separation part and the outer barrel. At least one hollow arm extends radially from the hub to the outer barrel, passing through the primary and secondary streams. A transmission shaft extends radially in the hollow arm. The hollow arm comprises a hydraulic-fluid outlet situated downstream of the transmission shaft. The arm further comprises a bypass channel or pocket able to bypass the transmission shaft. 1. An intermediate casing for a turbofan comprising a radially internal annular hub , a radially external annular barrel and an annular flow-separation part situated radially between the hub and the outer barrel , a primary stream for flow of a primary flow being delimited between the hub and the separation part , a secondary stream allowing flow of a secondary flow being delimited between the separation part and the outer barrel , at least one hollow arm extending radially from the hub to the outer barrel passing through the primary and secondary streams , a transmission shaft extending radially in said hollow arm , wherein the hollow arm comprises a hydraulic-fluid outlet situated downstream of the transmission shaft with respect to the direction of circulation of the primary flow or secondary flow , said arm further comprising a bypass channel or pocket able to bypass the transmission shaft and extending from upstream to downstream of said transmission shaft.2. An intermediate casing according to claim 1 , wherein the arm comprises first and second walls externally delimiting the arm claim 1 , extending radially and joining at an upstream edge claim 1 , said bypass channel or pocket being formed by a hollow region produced in the first wall and/or the second wall of the arm and ...

BLEED VALVE ASSEMBLY

A bleed valve assembly according to an exemplary aspect of the present disclosure includes, among other things, a bleed adaptor having an inlet portion, a fitting opposite the inlet portion, an adaptor body that extends between the inlet portion and the fitting, and a bleed opening disposed on the adaptor body that is selectively exposed to direct fluid into the bleed adaptor. 1. A bleed valve assembly , comprising: an inlet portion;', 'a fitting opposite said inlet portion;', 'an adaptor body that extends between said inlet portion and said fitting; and', 'a bleed opening disposed on said adaptor body that is selectively exposed to direct fluid into said bleed adaptor., 'a bleed adaptor having2. The bleed valve assembly as recited in claim 1 , wherein said fluid includes at least one of air claim 1 , mist and fuel.3. The bleed valve assembly as recited in claim 1 , comprising a hose connected to said fitting.4. The bleed valve assembly as recited in claim 1 , comprising a nut and a threaded portion between said inlet portion and said fitting.5. The bleed valve assembly as recited in claim 4 , comprising a seal between said nut and said threaded portion.6. The bleed valve assembly as recited in claim 1 , wherein an inlet portion of said bleed adaptor is received against a seat of a tube boss to prevent said fluid from entering said bleed adaptor.7. The bleed valve assembly as recited in claim 6 , wherein said bleed opening is disposed on said inlet portion.8. The bleed valve assembly as recited in claim 6 , wherein said inlet portion of said bleed adaptor is selectively spaced from said seat to direct said fluid into said bleed opening.9. The bleed valve assembly as recited in claim 1 , wherein an inlet portion of said bleed adaptor is moveable away from a seat of a tube boss to expose said bleed opening.10. The bleed valve assembly as recited in claim 1 , wherein said bleed adaptor is threadably received by a tube boss.11. A gas turbine engine claim 1 , comprising: ...

COOLING SYSTEM FOR GAS TURBINE, GAS TURBINE EQUIPMENT PROVIDED WITH SAME, AND PARTS COOLING METHOD FOR GAS TURBINE

A cooling system includes: a high pressure bleed line configured to bleed high pressure compressed air from a first bleed position of a compressor and to send the air to a first hot part; a low pressure bleed line configured to bleed low pressure compressed air from a second bleed position of the compressor and to send the air to a second hot part; an orifice provided in the low pressure bleed line; a connecting line configured to connect the high pressure bleed line and the low pressure bleed line; a first valve provided in the connecting line; a bypass line configured to connect the connecting line and the low pressure bleed line; and a second valve provided in the bypass line. 120-. (canceled)21. A cooling system for a gas turbine which includes a compressor configured to compress air , a combustor configured to burn a fuel in the air compressed by the compressor to generate a combustion gas , and a turbine driven using the combustion gas , the cooling system for a gas turbine comprising:a high pressure bleed line configured to bleed air from a first bleed position of the compressor and to send the air bled from the first bleed position to a first hot part coming into contact with the combustion gas among parts constituting the gas turbine;a cooler configured to cool air passing through the high pressure bleed line;a low pressure bleed line configured to bleed air at a pressure lower than that of the air which is bled from the first bleed position from a second bleed position of the compressor, to send the air bled from the second bleed position to a second hot part coming into contact with the combustion gas and disposed under a lower pressure environment than the first hot part among the parts constituting the gas turbine, and is not provided with a cooler;a minimum flow rate securing device configured to secure a minimum flow rate of air flowing through the low pressure bleed line while limiting a flow rate of the air flowing through the low pressure bleed ...

FEEDFORWARD CONTROL OF A FUEL SUPPLY CIRCUIT OF A TURBOMACHINE

A fuel supply system for a turbomachine, comprising a fuel circuit comprising pressurizer at the output of the circuit, a pump arranged to send into the circuit a fuel flow rate which is an increasing function of the rotational speed of a shaft of the pump, and a control circuit arranged to control the device to comply with a flow rate setpoint at the output of the fuel circuit. The system further comprises a feedforward corrector circuit configured to calculate an increment of the flow rate setpoint as a function of the engine speed of the turbomachine and of a variation in the engine speed of the turbomachine, and to add this increment to the flow rate setpoint. A method of regulating the pump is also described.

LOW FAN NOISE TURBOJET

A double flow turbojet includes a fan including a disk centered on an axis of the fan which is provided with fan blades on its periphery, the blades having a leading edge, and an air inlet sleeve extending upstream of the fan and configured to delimit a gas flow designed to enter into the fan the air inlet sleeve having a collecting surface, the turbojet having an aspect ratio 2. The turbojet according to claim 1 , wherein the form factor is comprised between 0.1 and 0.45.3. The turbojet according to claim 1 , wherein the turbojet has a bypass ratio greater than or equal to 10.4. The turbojet according to claim 1 , wherein the diameter of the fan is comprised between 203.2 centimeters and 279.4 centimeters.530. The turbojet according to claim 1 , wherein an upstream portion of the air inlet sleeve is not symmetrical.6. The turbojet according to claim 5 , wherein a downstream portion of the air inlet sleeve is axisymmetric claim 5 , a connection between the non-symmetrical upstream portion of the air inlet sleeve and its downstream axisymmetric portion extending at a distance comprised between one and five centimeters from a plane situated at the intersection between a radially internal wall of the air inlet sleeve and a most upstream point of the leading edge of the fan blades.7. The turbojet according to claim 1 , further comprising:a primary flow space and a concentric secondary flow space,a turbine, housed in the primary flow space and in fluid communication with the fan, anda reduction mechanism, coupling the turbine and the fan, said reduction mechanism comprising a star or planetary gear reduction mechanism having a reduction ratio comprised between 2.5 and 5.8. An aircraft comprising the turbojet according to .10. The method of dimensioning according to claim 9 , further comprising a step during which the form factor is defined so that said ratio is comprised between 0.1 and 0.45.11. The turbojet according to claim 2 , wherein the form factor is comprised ...

Aircraft engine fan

A gas turbine engine system has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The fan bypass inlet mass flow rate at the reference operating point is appreciably higher than the mass flow rate through the bypass duct at the peak bypass efficiency at a given fan reference rotational speed and cruise conditions. This results in increased design flexibility and improved overall engine performance.

HIGH EFFICIENCY GAS TURBINE ENGINE

A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox. 2. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 0.03 KgsNK≤Q≤0.035 KgsNK.3. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , 031 KgsNK≤Q≤0.034 KgsNK.4. A gas turbine engine according to claim 1 , wherein at cruise conditions claim 1 , the specific thrust is less than 100 Nkgs.5. A gas turbine engine according to claim 1 , wherein a fan tip loading is defined as dH/Utip claim 1 , where dH is the enthalpy rise across the fan and Utip is the translational velocity of the fan blades at the tip of the leading edge claim 1 , and at cruise conditions claim 1 , 0.28 JkgK/(ms) Подробнее

EXPANDING SHELL FLOW CONTROL DEVICE

A gas turbine engine includes a bypass flowpath between an outer engine case structure and a core engine. The bypass flow exits the engine through a nozzle. A flow control device that can expand or contract is arranged around the nozzle to control the bypass flow and includes a plurality of overlapping arcuate segments. A method of controlling a bypass flow includes providing a flow control device with overlapping segments that defines a bypass flow path, and actuating the segments to change the amount of overlap between segments and therefore the size of the bypass flow path.

TWO-PART VALVUE MEMBER ASSEMBLY

A two-part wastegate valve member assembly comprises a support member and a valve member. The support member defines an aperture. The valve member comprises a central portion extending through the aperture and two opposed end portions disposed on opposite sides of the aperture. Each of the two end portions has dimensions such that the valve member is held captive by the support member. The central portion and two opposed end portions of the valve member are integrally formed. A method for forming the two-part wastegate valve member assembly comprises casting a single manufacturing intermediate and subsequently processing the manufacturing intermediate so as to form the two-part assembly. 1. A method for forming a two-part wastegate valve member assembly , the method comprising:casting a single manufacturing intermediate;processing the manufacturing intermediate so as to form a two-part assembly, the two-part assembly comprising a support member and a valve member, wherein the support member defines an aperture and wherein the valve member comprises a central portion extending through the aperture and two opposed end portions disposed on opposite sides of the aperture each of the two end portions having dimensions such that the valve member is held captive by the support member.2. The method of wherein the casting of the single manufacturing intermediate is achieved using investment casting.3. The method of wherein casting the single manufacturing intermediate includes:forming first and second temporary patterns, each of the first and second temporary patterns corresponding to a different portion of the manufacturing intermediate, wherein the first temporary pattern comprises a portion that corresponds to a portion of the manufacturing intermediate from which part of the valve member will be formed and the second temporary pattern comprises a portion that corresponds to a portion of the manufacturing intermediate from which a second part of the valve member will be ...

METHOD OF SUPPLYING FUEL TO AN INTERNAL FUEL MANIFOLD

The described method of supplying fuel to an internal fuel manifold of a bypass gas turbine engine includes directing a fuel flow through a fuel fairing having an outer surface exposed to a cool bypass airflow, directing the fuel flow in the fuel fairing through a heat exchanging structure on the outer surface of the fuel fairing to cool the fuel flow being below a coking temperature of the fuel, and then feeding the cooled fuel flow from the fuel fairing to the internal fuel manifold. 1. A method of supplying fuel to an internal fuel manifold of a bypass gas turbine engine , the method comprising:directing a fuel flow through a fuel fairing having an outer surface exposed to a cool bypass airflow;directing the fuel flow in the fuel fairing through a heat exchanging structure on the outer surface of the fuel fairing to cool the fuel flow below a coking temperature of the fuel; and subsequentlyfeeding the cooled fuel flow from the fuel fairing to the internal fuel manifold.2. The method of claim 1 , wherein directing fuel through the fuel fairing includes directing fuel through a fuel channel extending through the fuel fairing between a fuel inlet and the fuel outlet claim 1 , the fuel channel having a portion disposed in a body of the fuel fairing directly behind the plurality of heat exchanging structures so as to be in direct conductive heat transfer communication therewith.3. The method of claim 2 , wherein directing fuel through the fuel channel includes directing fuel through a first portion of the fuel channel claim 2 , directing fuel through a U-turn connected to the first portion claim 2 , and directing fuel through a second portion of the fuel channel claim 2 , the first portion being the portion directly behind the plurality of heat exchanging structures.4. The method of claim 1 , wherein cooling the fuel in the fuel fairing below the predetermined temperature comprises using a plurality of heat exchanging fins disposed on the outer surface of the fuel ...

Combined Cycle Power Plant and Start-Up Method of the Same

There is provided a combined cycle power plant in which a high-pressure steam turbine and an intermediate-pressure steam turbine can operate in a state where amounts of thermal effect thereof are close to a limit value, and capable of reducing start-up time. A combined cycle power plant includes: an exhaust heat recovery boiler that includes a high-pressure superheater which superheats steam for a high-pressure steam turbine, and a reheater which reheats steam for an intermediate-pressure steam turbine; bypass pipes through which steam bypasses the high-pressure superheater and the reheater; bypass valves that regulate flow rates of steam which flows through the bypass pipes; and a bypass controller that controls the bypass valves such that a difference between thermal effect-amount margins of the turbines is decreased.

GAS TURBINE ENGINE WITH ROTOR TIP CLEARANCE CONTROL SYSTEM

A gas turbine engine includes a turbine and a rotor tip clearance control system. The rotor tip clearance control system is configured to actively manage a clearance formed between a rotor of the turbine and a case structure of the turbine. 1. A gas turbine engine , the engine comprisinga multi-stage axial compressor configured to compress air drawn into the engine and discharge pressurized air,a combustor configured to combust fuel in pressurized air from the compressor so as to create hot, high pressure combustion products,a turbine configured to receive the combustion products and to extract mechanical work from the combustion products as the combustion products move through the turbine, the turbine including a rotor with blades mounted for rotation about an axis and a case that extends around the rotor to block combustion products from moving though the turbine without interaction with the blades, anda rotor tip clearance control system configured to actively manage a clearance formed between the rotor and the case of the turbine, the rotor tip clearance control system including (i) a first flow modulator configured to control a cool-air flow from a first bleed location within the compressor so as to control the cool-air flow, (ii) a second flow modulator configured to control a warm-air flow from a second bleed location within the compressor, the warm-air flow being warmer than the cool-air flow and the second bleed location being downstream of the first bleed location, so as to control the warm-air flow, and (iii) an air temperature unit configured to receive the cool-air flow and the warm-air flow, the air temperature unit being configured to discharge a mixed-air flow made up of air from the cool-air flow and the warm-air flow to the case of the turbine in order to adjust a diameter of the case based on thermal expansion or contraction induced by the mixed-air flow,wherein the air temperature unit includes (a) a heat exchanger that conducts the cool-air flow ...

STATOR CONFIGURATION FOR GAS TURBINE ENGINE

A stator configuration for a gas turbine engine including a splitter segment. Also included is a first stator extending radially outwardly from the splitter segment. Further included is a second stator extending radially inwardly from the splitter segment, the splitter segment, the first stator and the second stator being a single, integrally formed structure. 1. A stator configuration for a gas turbine engine comprising:a splitter segment;a first stator extending radially outwardly from the splitter segment; anda second stator extending radially inwardly from the splitter segment, the splitter segment, the first stator and the second stator being a single, integrally formed structure.2. The stator configuration of claim 1 , wherein the splitter segment is configured to split an incoming fluid flow into a first claim 1 , radially outer stream and a second claim 1 , radially inner stream.3. The stator configuration of claim 1 , wherein the splitter segment extends from a forward end to a rearward end claim 1 , the first stator positioned closer to the forward end relative to the distance between the second stator and the forward end.4. The stator configuration of claim 1 , wherein the splitter segment claim 1 , the first stator and the second stator are formed from a single manufacturing process to form the single claim 1 , integrally formed structure.5. The stator configuration of claim 1 , wherein the splitter segment claim 1 , the first stator and the second stator are multiple components joined together to form the single claim 1 , integrally formed structure.6. The stator configuration of claim 1 , further comprising a flange extending radially inwardly from the second stator claim 1 , the flange providing a location for operative coupling of the stator configuration to a stationary structure of the gas turbine engine.7. The stator configuration of claim 6 , further comprising a protrusion extending rearward from the flange to provide a locating feature during ...

Intercooled cooling air with auxiliary compressor control

A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.

BELLMOUTH NOZZLE ASSEMBLY FOR A GAS TURBINE ENGINE

A gas turbine engine that includes a nozzle assembly that has features that facilitate airflow into and through a bypass passage of the gas turbine engine during a reverse thrust operation is provided. The nozzle assembly of the gas turbine engine also includes features that increase the effectiveness of the thrust reverse system of the gas turbine engine. Methods for reversing the thrust of a gas turbine engine are also provided. 1. A gas turbine engine defining an outlet and an axial direction , a radial direction , and a circumferential direction , the gas turbine engine comprising:a core turbine engine;a nacelle disposed about the core turbine engine along the circumferential direction, the nacelle extending between a first end and a second end along the axial direction; and an outer panel coupled with the nacelle, the outer panel movable along the radial direction to move the nozzle assembly between the stowed position and the deployed position; and', 'an elastic member coupled with the outer panel and with the nacelle, wherein when the nozzle assembly is in the deployed position, the elastic member is inflated with an airflow such that the elastic member forms a bellmouth at the outlet of the gas turbine engine., 'a nozzle assembly disposed at or proximate the second end of the nacelle and movable between a stowed position and a deployed position, the nozzle assembly comprising2. The gas turbine engine of claim 1 , wherein the nacelle is spaced from the core turbine engine along the radial direction so as to define a bypass passage therebetween claim 1 , and wherein the outlet is a bypass passage outlet.3. The gas turbine engine of claim 1 , further comprising:a thrust reverser system, wherein the thrust reverser system is a variable pitch fan assembly.4. The gas turbine engine of claim 1 , wherein the nacelle comprises an outer surface and wherein the nacelle defines a recess along the outer surface claim 1 , and wherein when the nozzle assembly is in the ...

SYSTEMS AND METHODS FOR CONTROLLING A BLEED-OFF VALVE OF A GAS TURBINE ENGINE

Methods and systems for controlling a bleed-off valve of a gas turbine engine are described. The method comprises maintaining a first bleed-off valve associated with a first compressor of the gas turbine engine at least partially open upon detection of an unintended engine disturbance causing a drop in pressure of a combustion chamber of the engine; monitoring a rotor acceleration of the first compressor; and controlling closure of the first bleed-off valve when the rotor acceleration of the first compressor reaches a first threshold for a first duration. 1. A method for controlling a bleed-off valve of a gas turbine engine , the method comprising:maintaining a first bleed-off valve associated with a first compressor of the gas turbine engine at least partially open upon detection of an unintended engine disturbance causing a drop in pressure of a combustion chamber of the engine;monitoring a rotor acceleration of the first compressor; andcontrolling closure of the first bleed-off valve when the rotor acceleration of the first compressor reaches a first threshold for a first duration.2. The method of claim 1 , wherein the first compressor is a low pressure compressor and the first bleed-off valve is a low pressure compressor bleed-off valve claim 1 , and wherein controlling closure of the first bleed-off valve comprises modulating the first bleed-off valve as a function of a speed of the low pressure compressor until closure.3. The method of claim 2 , further comprising monitoring a rotor acceleration of a high pressure compressor of the engine claim 2 , and wherein controlling closure of the first bleed-off valve comprises controlling closure when the rotor acceleration of the low pressure compressor reaches the first threshold for the first duration and the rotor acceleration of the high pressure compressor reaches a second threshold for a second duration.4. The method of claim 2 , further comprising disabling a nominal mode of control of the first bleed-off valve ...

INFLATABLE CASCADE ASSEMBLY, SYSTEM, AND METHOD FOR A CASCADE THRUST REVERSER SYSTEM

There is provided an inflatable cascade assembly for a cascade thrust reverser system of an engine of an air vehicle. The inflatable cascade assembly has inflatable cascade members for inflation with a pressurized fluid. The inflatable cascade members are movable between a stowed deflated state and a deployed inflated state. Each inflatable cascade member has a forward end, an aft end, and a body. The body has circumferential vanes each having a first non-inflatable rigid side attached adjacent to a second inflatable flexible side. The body further has inflatable support members that are spaced apart and longitudinally extending, and coupled in a perpendicular arrangement to the circumferential vanes. The body further has a plurality of flow openings defined between the circumferential vanes and the inflatable support members. Each inflatable cascade member further has first and second extendable side supports coupled to respective first and second side ends of the body. 1. An inflatable cascade assembly for a cascade thrust reverser system of an engine of an air vehicle , the inflatable cascade assembly comprising: a forward end;', 'an aft end;', a plurality of circumferential vanes that are spaced apart and laterally extending, each circumferential vane comprising a first non-inflatable rigid side attached adjacent to a second inflatable flexible side;', 'a plurality of inflatable support members that are spaced apart and longitudinally extending, the plurality of inflatable support members being coupled in a perpendicular arrangement to the plurality of circumferential vanes; and', 'a plurality of flow openings defined between the plurality of circumferential vanes and the plurality of inflatable support members;, 'a body formed between the forward end and the aft end, the body comprising, 'a first extendable side support coupled to first side ends of the body; and', 'a second extendable side support coupled to second side ends of the body, the first extendable ...

TURBOCHARGER

A turbocharger includes a turbine impeller, a turbine housing which includes a scroll portion and a discharge port, a bypassing passage portion which guides a working fluid from the scroll portion to the discharge port, and a valve portion which controls the inflow of the working fluid to the bypassing passage portion. The hub side scroll is formed so that a cross-sectional area of a passage is larger than that of a shroud side scroll. The valve portion includes a hub side valve which controls the inflow of the working fluid from the hub side scroll to the bypassing passage portion and a shroud side valve which controls the inflow of the working fluid from the shroud side scroll to the bypassing passage portion. An operation of opening and closing the hub side valve is independent from an operation of opening and closing the shroud side valve. 1. A turbocharger comprising:a turbine impeller which rotates by using a predetermined axis as a rotation axis;a housing which includes a scroll portion formed to surround the turbine impeller and supplying a working fluid to the turbine impeller and a discharge portion discharging the working fluid passing through the turbine impeller;a bypassing passage portion of which one end is connected to the scroll portion and the other end is connected to the discharge portion and which guides the working fluid from the scroll portion to the discharge portion; anda valve portion which is provided in the bypassing passage portion and controls an inflow of the working fluid from the scroll portion to the discharge portion,wherein the scroll portion includes a first scroll and a second scroll,wherein the first scroll is formed so that a cross-sectional area of a passage is larger than that of the second scroll,wherein the valve portion includes a first valve and a second valve,wherein the first valve controls the inflow of the working fluid from the first scroll to the bypassing passage portion,wherein the second valve controls the ...

TURBOMACHINERY TRANSITION DUCT FOR WIDE BYPASS RATIO RANGES

A gas turbine engine includes a case assembly, a splitter, an upstream blade row, and a transition duct. The case assembly defines an outer flow path wall and an inner flow path wall. The splitter is disposed between the outer flow path wall and the inner flow path wall. The splitter has a first surface and a second surface disposed opposite the first surface. The transition duct is defined by the outer flow path and the inner flow path and extends between the upstream blade row and the leading edge of the splitter. 1. A gas turbine engine comprising:a case assembly defining an outer flow path wall and an inner flow path wall, each extending between a first case end and a second case end along a central longitudinal axis;a splitter disposed between the outer flow path wall and the inner flow path wall, the splitter having a first surface and a second surface disposed opposite the first surface, each extending from the second case end towards a leading edge;an upstream blade row is disposed proximate the first case end and extends between the outer flow path wall and the inner flow path wall; anda transition duct defined by the outer flow path and the inner flow path and extends between the upstream blade row and the leading edge, the transition duct defining an entrance section that extends from the upstream blade row towards a diverging section that extends towards the leading edge.2. The gas turbine engine of claim 1 , wherein the entrance section has an entrance annular area.3. The gas turbine engine of claim 2 , wherein the transition duct includes a first outlet having a first annular area and a second outlet having a second annular area.4. The gas turbine engine of claim 3 , wherein a combination of the first annular area and the second annular area is greater than the entrance annular area.5. The gas turbine engine of claim 1 , wherein an outer duct is defined between the outer flow path wall and the first surface and extends from the leading edge towards the ...

MECHANICALLY DRIVEN AIR VEHICLE THERMAL MANAGEMENT DEVICE

The present disclosure is directed to an aircraft power generation system including a reverse Brayton cycle system, a gas turbine engine, and a gearbox. The gas turbine engine includes a compressor section, a turbine section, and an engine shaft. The compressor section is arranged in serial flow arrangement with the turbine section. The engine shaft is rotatable with at least a portion of the compressor section and with at least a portion of the turbine section. The reverse Brayton cycle system includes a compressor, a driveshaft, a turbine, and a first exchanger. The driveshaft is rotatable with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine are in serial flow arrangement. The gearbox is configured to receive mechanical energy from the engine shaft and transmit mechanical energy to the reverse Brayton cycle system through the driveshaft. 1. An aircraft power generation system , comprising:a gas turbine engine including a compressor section, a turbine section, and an engine shaft, the compressor section arranged in serial flow arrangement with the turbine section, and the engine shaft rotatable with at least a portion of the compressor section and with at least a portion of the turbine section;a reverse Brayton cycle system, including a compressor, a driveshaft, a turbine, and a first exchanger, the driveshaft rotatable with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine in serial flow arrangement; anda gearbox, wherein the gearbox is configured to receive mechanical energy from the engine shaft and transmit mechanical energy to the reverse Brayton cycle system through the driveshaft.2. The system in claim 1 , further comprising:a thermal management system; anda working fluid, wherein the working fluid is in the reverse Brayton cycle system, and wherein the working fluid is in fluid communication with the thermal management system.3. The system in claim 2 , wherein the ...

HEAT ENGINE AND METHOD FOR OPERATING A HEAT ENGINE

Heat engine and method for operating a heat engine are disclosed. A disclosed method for operating a heat engine-includes supplying gas containing combustible constituents as supply air to a combustion chamber of a combustion apparatus-of the heat engine; supplying fuel to the combustion chamber; removing exhaust gas from the combustion chamber and supplying the exhaust gas to a heat exchanger of the heat engine; transferring heat from the exhaust gas to at least a part of the supply air by the heat exchanger; guiding a part of at least one of the supply air or a part of the exhaust gas past the heat exchanger by a bypass guide, wherein at least one of a mass stream or volumetric stream of at least one of the part of the supply air or the exhaust gas guided past the heat exchanger-is at least one of controlled or regulated-so that at least one of a thermal power or a mechanical power of the heat engine is approximately constant over time. 1. Method for operating a heat engine , the method comprising:supplying gas containing combustible constituents as supply air to a combustion chamber of a combustion apparatus of the heat engine;supplying fuel to the combustion chamber;removing exhaust gas from the combustion chamber and supplying the exhaust gas to a heat exchanger of the heat engine;transferring heat from the exhaust gas to at least a part of the supply air by the heat exchanger; andguiding a part of at least one of the supply air or a part of the exhaust gas past the heat exchanger by a bypass guide, wherein at least one of a mass stream or volumetric stream of at least one of the part of the supply air or the exhaust gas guided past the heat exchanger is controlled so that at least one of a thermal power or a mechanical power of the heat engine is approximately constant over time.2. Method in accordance with claim 1 , wherein at least one of a mass stream or volumetric stream of the fuel supplied to the combustion chamber is approximately constant over time.3. ...

GAS TURBINE ENGINE MOUNTING ARRANGEMENT

A gas turbine engine () comprises a bypass duct cowl (), an engine core housing () defining an engine core inlet, a bypass fan () and a plurality of outlet guide vanes (). Each outlet guide vane extends between a radially inner surface of the bypass duct cowl () and a radially outer surface of the engine core housing () to define an outlet guide vane span (SPANOGV). The outlet guide vanes () are configured to support the engine core housing () relative to the bypass duct cowl (). The bypass fan () and an engine core inlet () define a bypass ratio between 10 and 17, and a ratio of the outlet guide vane span (OGVSPAN) to a bypass fan radius (RFAN) is between 0.45 and 0.55. 1. A gas turbine engine comprising:a bypass duct cowl;an engine core housing defining an engine core inlet;a bypass fan; anda plurality of outlet guide vanes extending between a radially inner surface of the bypass duct cowl, and a radially outer surface of the engine core housing to define an outlet guide vane span, the outlet guide vanes being configured to support the engine core housing relative to the bypass duct cowl;wherein the bypass fan and an engine core inlet define a bypass ratio between 10 and 17; and a ratio of the outlet guide vane span to a bypass fan radius is between 0.45 and 0.55.3. A gas turbine engine according to claim 1 , wherein a ratio of the inner radius of the outlet guide vanes and an outer radius of the outlet guide vanes is between 0.4 and 0.6.4. A gas turbine engine according to claim 3 , wherein the ratio of the inner radius of the outlet guide vanes and the outer radius of the outlet guide vanes is between 0.5 and 0.55.5. A gas turbine engine according to claim 1 , wherein a ratio of an axial distance between an inlet to the bypass duct cowl and a mid chord of the outlet guide vanes at the tip claim 1 , and an outer radius of the outlet guide vanes is between 1 and 1.8.6. A gas turbine according to claim 5 , wherein the ratio of an axial distance between an inlet to ...

Conduit cooling system and method of supplying cooling fluid to a conduit

A conduit system for a gas turbine engine includes, a heat exchanger configured to cool fluid flowing therethrough having an inlet and an outlet, at least one by-pass in operable communication with the heat exchanger that is configured to allow fluid to exit the heat exchanger before reaching the outlet, and a conduit that is in fluidic communication with the outlet and the at least one by-pass.

Combustion bypass passive valve system for a gas turbine

A combustor for a gas turbine, including: a combustor chamber; a casing enclosing the combustor chamber and defining an area therebetween for passing compressor discharge air into the combustor chamber for use in combustion; and at least one passive bypass valve for selectively extracting a portion of the compressor discharge air from the area between the combustor chamber and the casing to adjust a temperature in the combustor.

GAS TURBINE ENGINE WITH AIRFLOW MEASUREMENT SYSTEM

A turbofan gas turbine engine having a bypass duct, and a bypass airflow measurement system. The bypass airflow measurement system comprises: at least one acoustic transmitter configured to transmit an acoustic waveform across the bypass duct of the gas turbine engine though which a bypass airflow passes to at least one acoustic receiver; where the at least one acoustic transmitter and the at least one acoustic receiver are located on an axial plane that is substantially perpendicular to the bypass flow. A method of measuring bypass airflow properties of a turbofan gas turbine engine is also described. 1. A turbofan gas turbine engine having a bypass duct and a bypass airflow measurement system , the bypass airflow measurement system comprising:at least one acoustic transmitter configured to transmit an acoustic waveform across the bypass duct of the gas turbine engine though which a bypass airflow passes to at least one acoustic receiver;where the at least one acoustic transmitter and the at least one acoustic receiver are located on an axial plane that is substantially perpendicular to the bypass flow.2. The turbofan gas turbine engine of claim 1 , wherein the axial plane upon which the at least one acoustic transmitter and the at least one acoustic receiver are located extends from 80° to 100° to the bypass flow.3. The turbofan gas turbine engine of claim 2 , wherein the axial plane upon which the at least one acoustic transmitter and the at least one acoustic receiver are located extends from 85° to 95° to the bypass flow.4. The turbofan gas turbine engine of claim 3 , wherein the axial plane upon which the at least acoustic transmitter and the at least one acoustic receiver are located extends about 90° to the bypass flow.5. The turbofan gas turbine engine of claim 1 , where the at least one acoustic transmitter and the at least one acoustic receiver are mounted in a casing that defines the bypass duct so that they do not substantially protrude into the bypass ...

GAS TURBINE ENGINE HAVING RADIALLY-SPLIT INLET GUIDE VANES

An apparatus for the control of fluid flow in a gas turbine engine comprises a first plurality of inlet guide vanes disposed upstream of a fan, a compressor, a combustor, and a turbine; at least one airflow splitter adapted to split air admitted through the first plurality of inlet guide vanes into a core airflow which flows through the fan, the compressor, the combustor, and the turbine and a bypass airflow which flows through the fan; wherein the first plurality of inlet guide vanes comprise a radially-inward first portion adapted to direct air admitted through the first plurality of inlet guide vanes to the core airflow and a radially-outward second portion adapted to direct air admitted through the first plurality of inlet guide vanes to the bypass airflow, and wherein the first portion comprises a fixed vane and the second portion comprises a variable vane. 1. An apparatus for the control of fluid flow in a gas turbine engine comprising:a first plurality of inlet guide vanes disposed upstream of a fan, a compressor, a combustor, and a turbine;at least one airflow splitter adapted to split air admitted through said first plurality of inlet guide vanes into a core airflow which flows through said fan, said compressor, said combustor, and said turbine and a bypass airflow which flows through said fan;wherein said first plurality of inlet guide vanes comprise a radially-inward first portion adapted to direct air admitted through said first plurality of inlet guide vanes to said core airflow and a radially-outward second portion adapted to direct air admitted through said first plurality of inlet guide vanes to said bypass airflow, and wherein said first portion comprises a fixed vane and said second portion comprises a variable vane.2. The apparatus of further comprising an actuator adapted to adjust the position of said second portion.3. The apparatus of wherein said variable vane comprises a fixed strut and a rotatable flap claim 2 , and wherein the orientation ...

System and Method for A Fluidic Barrier From the Upstream Splitter

A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of high pressure fluid jets originating from a trailing edge of the upstream splitter, the jets restricting the migration of the core stream into the bypass stream through the fluid passage. 1. A turbofan engine having a fan portion in fluid communication with a core stream and a bypass stream of air; the core stream being:compressed by the fan portion and a core compressor portion, heated and expanded through a core turbine portion;the core turbine portion driving the fan and the compressor portion; the core turbine portion connected to a shaft;the bypass stream being compressed by the fan portion;the core and the bypass streams separated by an upstream splitter and a downstream splitter with the fan portion disposed axially between the upstream and downstream splitters wherein a fluid passage between the core and bypass streams is defined between the splitters and blades of the fan portion; anda plurality of high pressure fluid jets originating from a trailing edge of the upstream splitter restricting the migration of the core stream into the bypass stream through the fluid passage.2. The turbofan engine of claim 1 , wherein the high pressure fluid jets originate from orifices on the trailing edge of the upstream splitter.3. The turbofan engine of claim 2 , further comprising a plurality of passages supplying high pressure compressed air to the fluid jets.4. The turbofan engine of claim 3 , wherein the plurality of passages are in fluid communication with the core compressor portion.5. The turbofan engine of claim 1 , wherein the pressure of the core stream is greater than the pressure of the bypass stream.6. The turbofan engine of wherein the plurality of orifices are distributed around the ...

INTERCOOLED COOLING AIR WITH DUAL PASS HEAT EXCHANGER

A gas turbine engine comprises a main compressor section having a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses ng air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed. 120.-. (canceled)21. An intercooling system for a gas turbine engine comprising:a heat exchanger for cooling air drawn from a portion of a main compressor section at a first temperature and pressure for cooling the air to a second temperature cooler than the first temperature;a cooling compressor compressing air communicated from the heat exchanger to a second pressure greater than the first pressure and communicating the cooling air to a portion of at least a turbine section; andsaid heat exchanger having at least two passes, with a first of said passes passing air in a direction having at least a radially outward component, and a second of said passes returning the air in a direction having at least a radially inward component to the compressor.22. The intercooling system as set forth in claim 21 , wherein said first pass is positioned downstream of said second pass in said bypass duct.23. The intercooling system as set forth in claim 22 , wherein said cooling compressor includes a centrifugal compressor impeller.24. The intercooling system as set forth in claim 22 , wherein a main fan delivers bypass air into a bypass duct and into said main compressor section and said heat exchanger positioned within said bypass duct to be cooled by bypass air.25. The intercooling ...

TURBINE ENGINE SYSTEM WITH HEAT EXCHANGER IN BYPASSABLE SECONDARY DUCT

An assembly is provided for a turbine engine. This assembly includes a primary duct, a bleed duct, a plurality of secondary ducts, a heat exchanger and a flow regulator. The bleed duct extends from a bleed duct inlet to a bleed duct outlet. The bleed duct inlet is fluidly coupled with the primary duct. The secondary ducts are arranged in parallel between the bleed duct outlet and the primary duct. The secondary ducts include a first duct and a second duct. The heat exchanger is configured with the second duct. The flow regulator is configured to direct at least a majority of fluid flowing through the bleed duct outlet to: (A) the first duct during a first mode of operation; and (B) the second duct during a second mode of operation. 1. An assembly for a turbine engine , comprising:a primary duct;a bleed duct extending from a bleed duct inlet to a bleed duct outlet, the bleed duct inlet fluidly coupled with the primary duct;a plurality of secondary ducts arranged in parallel between the bleed duct outlet and the primary duct, the plurality of secondary ducts including a first duct and a second duct;a heat exchanger configured with the second duct; and the first duct during a first mode of operation; and', 'the second duct during a second mode of operation., 'a flow regulator configured to direct at least a majority of fluid flowing through the bleed duct outlet to'}2. The assembly of claim 1 , wherein the flow regulator is configured to direct at least substantially all of the fluid flowing through the bleed duct outlet tothe first duct during the first mode of operation; andthe second duct during the second mode of operation.3. The assembly of claim 1 , wherein the flow regulator comprises a two way valve.4. The assembly of claim 1 , wherein the flow regulator is located between the bleed duct and the first duct.5. The assembly of claim 1 , wherein the flow regulator is located between the bleed duct and the second duct.6. The assembly of claim 1 , whereinthe flow ...

EXHAUST GAS GUIDE SECTION FOR AN EXHAUST GAS TURBOCHARGER AND METHOD FOR OPERATING AN EXHAUST GAS TURBOCHARGER

The invention relates to an exhaust gas guide section for an exhaust gas turbocharger, with a control device, comprising a cover element () for opening and closing a bypass duct () of the exhaust gas guide section (), wherein the bypass duct () is provided in the through-flow exhaust gas guide section () for bypassing a turbine wheel which is rotatably accommodated in a wheel chamber of the exhaust gas guide section (), and wherein the bypass duct () comprises a second flow cross section (). 119311. An exhaust gas guide section () for an exhaust gas turbocharger , with a control device , comprising a cover element () for opening and closing a bypass duct () of the exhaust gas guide section () ,{'b': 31', '1', '1, 'wherein the bypass duct () is provided in a through-flow exhaust gas guide section () for bypassing a turbine wheel which is rotatably accommodated in a wheel chamber of the exhaust gas guide section (), and'}{'b': 31', '33, 'claim-text': {'b': 33', '9, 'wherein the second flow cross section () may be degressively opened by means of the cover element () at least starting from a closed position into an intermediate position.'}, 'wherein the bypass duct () comprises a second flow cross section (), and'}2. The exhaust gas guide section according to claim 1 ,{'b': 4', '5', '1, 'wherein a first spiral channel () and a second spiral channel () for the inflow of a turbine wheel are formed in the exhaust gas guide section (),'}{'b': 10', '4', '5', '4', '5, 'wherein a through-flow opening () is provided between the first spiral channel () and the second spiral channel (), which is configured for inducing an overflow of exhaust gas from the first spiral channel () into the second spiral channel () and vice versa, and'}{'b': 10', '9', '10, 'wherein the through-flow opening () comprises a movable covering element (′) for opening or closing of the through-flow opening ().'}3. The exhaust gas guide section according to claim 2 ,{'b': 9', '11', '10, 'wherein the covering ...

Heat exchange systems for turbomachines

The disclosure concerns a heat exchange system or cooling system for a flow machine ( 10 ) having a cooling duct ( 30 ) with a coolant inlet opening ( 32 ) and a closure ( 34 ) for selectively opening the inlet opening ( 32 ). A component ( 38 ) is arranged to be fluid washed by flow along the cooling duct ( 30 ). A flow injector ( 40 ) is spaced from the closure ( 34 ) along the cooling duct ( 30 ) and oriented to inject flow into the cooling duct ( 30 ) in a direction that creates a negative fluid pressure downstream of the closure ( 34 ), wherein the closure ( 34 ) is openable in response to said negative pressure. The component may be a heat exchanger ( 38 ). The flow injector ( 40 ) may be fed by a compressor ( 14, 15 ) of the flow machine ( 10 ).

TURBOMACHINE AND SYSTEM FOR COMPRESSOR OPERATION

A turbomachine defining a flowpath therethrough at which a fluid is compressed is provided, in which the turbomachine includes a first compressor in serial flow arrangement upstream of a second compressor. The second compressor includes a port at the flowpath and is configured to receive at least a portion of the fluid from the flowpath from the second compressor. The first compressor includes a vane positioned at the flowpath and the vane includes an opening at the flowpath configured to egress the portion of the fluid from the port into the flowpath. 1. A turbomachine , wherein the turbomachine defines a flowpath therethrough at which a fluid is compressed , the turbomachine comprising:a first compressor in serial flow arrangement upstream of a second compressor,wherein the second compressor comprises a port at the flowpath, wherein the port at the second compressor is configured to receive at least a portion of the fluid from the flowpath from the second compressor, andwherein the first compressor comprises a vane positioned at the flowpath, and wherein the vane comprises an opening at the flowpath, wherein the opening at the vane is configured to egress the portion of the fluid from the port into the flowpath.2. The turbomachine of claim 1 , wherein the port at the second compressor is positioned at or aft of a downstream end of the second compressor.3. The turbomachine of claim 2 , wherein the opening at the vane of the first compressor is positioned between a maximum thickness location at the vane and 10% of a chord from a trailing edge of the vane.4. The turbomachine of claim 3 , the turbomachine comprising:a conduit extended from the port at the second compressor to the vane at the first compressor, wherein the conduit provides fluid communication from the flowpath at the downstream end of the second compressor, or aft thereof, to the opening at the vane.5. The turbomachine of claim 3 , wherein the vane comprises a plurality of the opening comprising a first ...

DISCHARGE FLOW DUCT OF A TURBINE ENGINE COMPRISING A VBV GRATING WITH VARIABLE SETTING

A hub of an intermediate casing for a dual-flow turbine engine includes a discharge flow duct extending between an inner shroud and an outer shroud of the hub, the discharge flow duct leading into the secondary flow space through an outlet opening formed in the outer shroud, the outlet opening included in a discharge plane substantially tangential to the outer shroud; and discharge fins including an upstream fin and a downstream fin. An upstream acute angle between the discharge plane and the skeleton line of the upstream fin is smaller than a downstream acute angle between the discharge plane and the skeleton line of the downstream fin. 1. A hub of an intermediate casing for a dual-flow turbine , said hub comprising:an inner shroud configured to delimit externally a primary flow space of a primary gas flow of the turbine engine,an outer shroud configured to delimit internally a secondary flow space of a secondary gas flow of said turbine engine,at least one discharge stream duct extending between the inner shroud and the outer shroud, the discharge stream duct leading into the secondary flow space through an outlet opening formed in the outer shroud, said outlet opening being comprised in a discharge plane substantially tangent to the outer shroud, anddischarge fins, attached in the discharge stream duct at the outlet opening of the outer shroud,said discharge fins comprising, from upstream to downstream in the gas flow direction in the secondary flow space, an upstream fin, extending adjacently to an upstream wall of the discharge stream duct and comprising a leading edge arranged facing a gas flow in the discharge stream duct and a skeleton line, and a downstream fin extending adjacently to a downstream wall of the discharge stream duct and comprising a leading edge and a skeleton line,{'sub': upstream', 'downstream, 'an upstream acute angle (α) between the discharge plane and a tangent to the skeleton line at the leading edge of the upstream fin being smaller ...

Variable bypass ratio fan with variable pitch aft stage rotor blading

A gas turbine engine includes a fan section. A splitter is downstream of the fan section and at least partially defines a secondary flow path on a radially outer side and an inner flow path on a radially inner side. A variable pitch rotor blade assembly is located at an inlet to the inner flow path and includes a plurality of variable pitch rotor blades. 1. A gas turbine engine comprising:a fan section;a splitter downstream of the fan section at least partially defining a secondary flow path on a radially outer side and an inner flow path on a radially inner side; anda variable pitch rotor blade assembly located at an inlet to the inner flow path including a plurality of variable pitch rotor blades.2. The gas turbine engine of claim 1 , wherein the plurality of variable pitch rotor blades rotate about a rotor blade axis that is transverse to an axis of rotation of the gas turbine engine.3. The gas turbine engine of claim 2 , wherein the rotor blade axis is perpendicular to an axis of rotation of the gas turbine engine.4. The gas turbine engine of claim 1 , wherein the fan section includes more than one fan blade row.5. The gas turbine engine of claim 1 , wherein the fan section includes at least one fan blade row with a stator row immediately downstream of the at least one fan blade row and immediately upstream of the variable pitch rotor blade assembly and the stator row includes a plurality of non-rotatable vanes.6. The gas turbine engine of claim 1 , wherein the fan section includes at least one fan blade row with a stator row immediately downstream of the at least one fan blade row and immediately upstream of the variable pitch rotor blade assembly and the stator row includes a plurality of rotatable vanes configured to rotate about an axis through a corresponding vane.7. The gas turbine engine of claim 1 , wherein the fan section includes a vane row immediately downstream of a fan blade row and immediately upstream of the inner flow path.8. The gas turbine ...

SUPERSONIC GAS TURBINE ENGINE

A supersonic gas turbine engine for an aircraft that comprises a nacelle, a fan, an engine core including a primary duct configured to guide a core airflow through the engine core, a bypass duct extending between the engine core and an engine casing and configured to guide a bypass airflow through the bypass duct, an intake located upstream of the fan, and a tertiary airflow duct extending between the engine casing and the nacelle and configured to guide a tertiary airflow. The intake is configured to extract air from the intake and guide it to the tertiary airflow duct in which the extracted air flows as tertiary airflow. It is provided that at least one heat exchanger is mounted in the tertiary airflow duct. 2. The gas turbine engine of claim 1 , wherein the at least one heat exchanger fills the entire annulus of the tertiary airflow duct in the circumferential direction.3. The gas turbine engine of claim 1 , wherein the at least one heat exchanger only partly fills the annulus of the tertiary airflow duct in the circumferential direction.4. The gas turbine engine of claim 3 , wherein a plurality of heat exchangers are mounted in the tertiary airflow duct claim 3 , wherein the heat exchangers are arranged to have gaps between them in the circumferential direction.5. The gas turbine engine of claim 1 , wherein at least one arc shaped heat exchanger is provided in the tertiary airflow duct.6. The gas turbine engine of claim 1 , wherein the at least one heat exchanger fills the entire annulus of the tertiary airflow duct in the radial direction.7. The gas turbine engine of claim 1 , wherein the at least one heat exchanger only partly fills the annulus of the tertiary airflow duct in the radial direction.8. The gas turbine engine of claim 7 , wherein the at least one heat exchanger is mounted to an inner wall of the nacelle.9. The gas turbine engine of claim 7 , wherein the at least one heat exchanger is mounted to an outer wall of the engine casing.10. The gas ...

MOVEABLE EXHAUST PLUG

An exhaust section of a gas turbine engine includes an exhaust plug defining a plurality of exhaust plug apertures circumferentially spaced from each other. Also included is an exhaust nozzle radially offset from the exhaust plug defining an exhaust pathway between the exhaust plug and the exhaust nozzle. Further included is an exhaust plug liner having a non-uniform outer surface axially aligned with the exhaust plug apertures. The exhaust plug liner is rotatable relative to exhaust plug between a first position and a second position to selectively change a cross-sectional area of the exhaust pathway during thrust reversal operation to increase an amount of reverse thrust. 1. An exhaust section of a gas turbine engine , comprising:an exhaust plug defining a plurality of exhaust plug apertures circumferentially spaced from each other;an exhaust nozzle radially offset from the exhaust plug defining an exhaust pathway between the exhaust plug and the exhaust nozzle; andan exhaust plug liner having a non-uniform outer surface axially aligned with the exhaust plug apertures;wherein the exhaust plug liner is rotatable relative to exhaust plug between a first position and a second position to selectively change a cross-sectional area of the exhaust pathway during thrust reversal operation to increase an amount of reverse thrust.2. The exhaust section of claim 1 , wherein the non-uniform outer surface of the exhaust plug liner is disposed along a single axial location of the exhaust plug liner.3. The exhaust section of claim 2 , wherein the non-uniform surface comprises at least one of a plurality of protrusions and recesses along the outer surface of the exhaust plug liner.4. The exhaust section of claim 1 , wherein the exhaust plug includes an end portion tapering to a plug tip.5. The exhaust section of claim 4 , wherein the end portion is conical.6. The exhaust section of claim 4 , further comprising an axially-extending portion of the exhaust plug extending from the ...

GAS TURBINE ENGINE WITH MINIMAL TOLERANCE BETWEEN THE FAN AND THE FAN CASING

A turbofan having a dilution rate of at least 10 and including a fan having a disc provided with blades at the periphery thereof, a distance between the head of the blades and the housing of the fan being less than or equal to ten millimeters; a primary flow space and secondary flow space that are concentric; a turbine, housed in the primary flow space and in fluid communication with the fan; and a reduction mechanism coupling the turbine and the fan. 1. A bypass gas turbine engine comprising:a fan housed in a fan casing, said fan comprising a disc provided with fan blades at its periphery, each fan blade comprising a blade tip extending at a distance from the disc,a primary flow space and a secondary concentric flow space,a turbine, housed in the primary flow space and in fluid communication with the fan, anda reduction mechanism, coupling the turbine and the fan,the gas turbine engine having a bypass ratio greater than or equal to 10, anda distance between the tip of the fan blades and the fan casing being less than or equal to ten millimeters.2. The gas turbine engine according to claim 1 , having a bypass ratio between 12 and 18.3. The gas turbine engine according to claim 1 , wherein the distance between the tip of the fan blades and the fan casing is less than or equal to six millimeters.4. The gas turbine engine according to claim 1 , wherein a thickness of the fan casing is less than or equal to fifteen millimeters.5. The gas turbine engine according to claim 1 , wherein an external diameter of the fan is between eighty inches and one hundred inches.6. The gas turbine engine according to claim 1 , wherein a difference in thickness of the fan casing claim 1 , between an upstream end and a downstream end of said fan casing claim 1 , is less than or equal to ten millimeters.7. The gas turbine engine according to claim 1 , wherein the fan casing is made of a composite material comprising a fibrous reinforcement densified by a matrix claim 1 , said fibrous ...

AIRCRAFT AND METHOD OF OPERATING SAME

The aircraft can have a first engine secured to a first wing on a first side of a fuselage, and a second engine secured to a second wing on a second side of the fuselage, the second wing having a proximal end secured to the fuselage, and a distal end extending away from the fuselage. While operating the first engine, compressed gas can be conveyed from the first engine to a thrust generating device located at the distal end of the second wing.

COMBUSTION ENGINE INCLUDING TURBOMACHINE

A combustion engine including at least one combustion chamber, a first bleed air supply fluidly coupled to a portion of the combustion engine upstream the combustion chamber, a second bleed air supply fluidly coupled to a portion of the combustion engine downstream the combustion chamber, a first thermal bus, and a turbomachine including a compressor, a rotary pump, and a first turbine, with the compressor and rotary pump in serial flow arrangement and the rotary pump being fluidly coupled to the first thermal bus. 1. A combustion engine comprising:at least one combustion chamber;a first bleed air supply fluidly coupled to a portion of the combustion engine upstream of the combustion chamber;a second bleed air supply fluidly coupled to a portion of the combustion engine downstream of the combustion chamber;a first thermal bus including a thermal source and a thermal dump fluidly coupled to one another and through which a heat transfer fluid flows; and a compressor, a rotary pump, and a first turbine mounted to a common shaft, with the compressor and the rotary pump in serial flow arrangement;', 'wherein the compressor is fluidly coupled to the first bleed air supply, the first turbine is fluidly coupled to the second bleed air supply, and the rotary pump is fluidly coupled to the first thermal bus to pump a heat transfer fluid through the first thermal bus from the thermal source to the thermal dump., 'a turbomachine comprising2. The combustion engine of claim 1 , wherein the turbomachine is fully integrated within the combustion engine.3. The combustion engine of claim 2 , wherein the combustion engine further comprises a housing and the turbomachine is located within an interior defined by the housing.4. The combustion engine of claim 1 , wherein at least one of the thermal source or the thermal dump is located within the combustion engine.5. The combustion engine of claim 4 , wherein the thermal source is fluidly coupled to at least a portion of an external ...

COMBUSTION ENGINE INCLUDING TURBOMACHINE

A combustion engine including at least one combustion chamber, a first bleed air supply fluidly coupled to a portion of the combustion engine upstream the combustion chamber, a second bleed air supply fluidly coupled to a portion of the combustion engine downstream the combustion chamber, a first thermal bus, and a turbomachine including a compressor, a rotary pump, and a first turbine, with the compressor and rotary pump in serial flow arrangement and the rotary pump being fluidly coupled to the first thermal bus. 1. A combustion engine comprising:at least one combustion chamber;a first bleed air supply fluidly coupled to a portion of the combustion engine upstream of the combustion chamber;a second bleed air supply fluidly coupled to a portion of the combustion engine downstream of the combustion chamber;a thermal bus including a thermal source and a thermal dump fluidly coupled to one another, with the thermal dump being downstream the thermal source, and through which a heat transfer fluid flows; and a compressor, a first turbine, and a second turbine mounted to a common shaft, with the first and second turbines in serial flow arrangement;', 'wherein the compressor is fluidly coupled to the first bleed air supply, the first turbine is fluidly coupled to the second bleed air supply, and the second turbine is fluidly coupled to the thermal bus, and wherein the heat transfer fluid is an engine exhaust gas from the second bleed air supply that flows through at least one of the first turbine or the second turbine., 'a turbomachine comprising a compressor;'}2. The combustion engine of claim 1 , wherein the thermal source is fluidly coupled to a subsystem of the combustion engine.3. The combustion engine of claim 2 , wherein the subsystem comprises at least one of a fuel system or an oil system for the combustion engine.4. The combustion engine of claim 1 , further comprising a fluid bypass between the first turbine and the second turbine such that at least a portion ...

BLEED VALVE ASSEMBLY FOR A GAS TURBINE ENGINE

A gas turbine engine includes a casing surrounding a first compressor, as well as a liner extending forward from the first compressor. Additionally, the gas turbine engine includes a bleed air assembly having a plurality of bleed valves positioned in the liner and spaced along a circumferential direction of the gas turbine engine. The bleed air assembly also includes a duct in airflow communication with the plurality of bleed valves and defining an outlet at the casing. The duct provides a flow bleed air from the plurality bleed valves to the outlet during operation. 1. A gas turbine engine defining a circumferential direction and comprising:a first compressor;a casing surrounding the first compressor;a liner extending forward from the first compressor; and a plurality of bleed valves positioned in the liner and spaced along the circumferential direction; and', 'a duct in airflow communication with the plurality of bleed valves and defining an outlet at the casing, the duct providing a flow of bleed air from the plurality of bleed valves to the outlet., 'a bleed air assembly comprising'}2. The gas turbine engine of claim 1 , wherein the duct extends continuously along the circumferential direction around the plurality of bleed valves.3360. The gas turbine engine of claim 2 , wherein the duct is an annular duct extending continuously three hundred and sixty degrees)(° along the circumferential direction.4. The gas turbine engine of claim 1 , wherein the gas turbine engine further defines a radial direction claim 1 , wherein the duct defines a first thickness along the radial direction proximate the outlet claim 1 , wherein the duct defines a second thickness along the radial direction at a location away from the outlet claim 1 , and wherein the first thickness is greater than the second thickness.5. The gas turbine engine of claim 4 , wherein the duct comprises a scroll liner claim 4 , wherein the first thickness is defined between the liner of the gas turbine engine ...

BLEED VALVE ASSEMBLY FOR A GAS TURBINE ENGINE

A gas turbine engine includes a first compressor, a casing surrounding the first compressor, and a liner extending forward from the first compressor. The gas turbine engine also includes a core turbine frame assembly extending between the liner and the casing and a bleed air assembly. The bleed air assembly includes a bleed valve positioned in the liner, and a duct in airflow communication with the bleed valve and defining in outlet. The duct is positioned within the core turbine frame assembly and extends to the casing. 1. A gas turbine engine defining an axial direction , the gas turbine engine comprising:a first compressor;a casing surrounding the first compressor;a liner extending forward from the first compressor;a core turbine frame assembly extending between the liner and the casing; and a bleed valve positioned in the liner; and', 'a duct in airflow communication with the bleed valve and defining an outlet, the duct positioned within the core turbine frame assembly and extending to the casing., 'a bleed air assembly comprising'}2. The gas turbine engine of claim 1 , wherein the core turbine frame assembly comprises a forward member and an aft member claim 1 , and wherein the duct is positioned between the forward member and the aft member.3. The gas turbine engine of claim 1 , wherein the outlet is positioned at the casing at a location forward of the aft member of the core turbine frame assembly along the axial direction and aft of the forward member of the core turbine frame assembly along the axial direction.4. The gas turbine engine of claim 1 , wherein the bleed valve comprises a door assembly movable between an open position and a closed position claim 1 , wherein the door assembly comprises a transition passage claim 1 , and wherein the transition passage is aligned with the duct when the door assembly is in the open position.5. The gas turbine engine of claim 4 , wherein the transition passage is misaligned with the duct when the door assembly is in ...

BYPASS VALVE ASSEMBLY FOR TURBINE GENERATORS

A bypass valve assembly for use in turbine generators includes a valve body defining a central bore and a plurality of passageways. Each passageway has a smaller area at an inlet portion and a larger area at an outlet portion to define a flared passageway. A plurality of bypass valves is disposed within the plurality of passageways within the valve body. Each bypass valve includes a base portion and a nose portion, with each nose portion defining a predefined contoured surface area. At least a portion of the contoured surface area includes a wear coating disposed thereon. Optionally, the wear coating includes a plasma enhanced magnetron sputtering nanocoating. 1. A bypass valve assembly for use in turbine generators comprising:a valve body defining a central bore and a plurality of passageways, each passageway having a smaller area at an inlet portion and a larger area at an outlet portion to define a flared passageway;a plurality of bypass seats disposed within each of the inlet portions of the passageways, the bypass seats being formed of a material having higher wear resistance than the valve body;a valve stem disposed within the central bore of the valve body;a valve cap secured to a distal end portion of the valve body;a bypass valve disc secured to a distal end portion of the valve stem;a plurality of bypass valves disposed within the plurality of passageways within the valve body, each bypass valve having a base portion and a nose portion, each nose portion defining a predefined contoured surface area, and at least a portion of the contoured surface area having a wear coating disposed thereon; anda pressure seal head disposed around a distal end portion of the valve stem, the pressure seal head defining proximal facing steps having a wear coating disposed thereon.2. The bypass valve assembly according to further comprising:at least one socket bolt securing the valve cap to the valve body; anda cap pin disposed under a head of the socket bolt and extending ...

BYPASS HOUSING IN AIR CYCLE MACHINE

A turbine bypass housing is provided. The housing includes a body having an inlet end and an outlet end and defining an interior, an outlet flange is configured at the outlet end and has a rim diameter, and an inlet flange is configured at the inlet end and has a flange thickness. An interior wall of the body has a first fillet radius proximal to the inlet flange and at least one boss is configured on an exterior of the body, wherein the boss defines a boss aperture therein, the boss aperture extending from an exterior of the body to the interior of the body. The flange thickness is 0.200±0.005 inches (0.508±0.013 cm). 1. A turbine bypass housing comprising:a body having an inlet end and an outlet end and defining an interior;an outlet flange is configured at the outlet end and has a rim diameter;an inlet flange is configured at the inlet end and has a flange thickness;an interior wall of the body has a first fillet radius proximal to the inlet flange;at least one boss is configured on an exterior of the body, wherein the boss defines a boss aperture therein, the boss aperture extending from an exterior of the body to the interior of the body,wherein the flange thickness is 0.200±0.005 inches (0.508±0.013 cm).2. The bypass housing of claim 1 , wherein the body has a general wall thickness of 0.095±0.020 inches (0.241±0.051 cm).3. The bypass housing of claim 1 , wherein the first fillet radius is 0.250±0.030 inches (0.635±0.076 cm).4. The bypass housing of claim 1 , wherein the body includes a general fillet radius claim 1 , and wherein a ratio between the first fillet radius and the general fillet radius is 1.330.5. The bypass housing of claim 1 , wherein the body has an interior radial length claim 1 , and the at least one boss defines a boss radial length claim 1 , wherein a ratio of the boss radial length to the interior radial length is 1.176.6. The bypass housing of claim 5 , wherein the interior radial length is 4.890±0.045 inches (12.421±0.114 cm).7. The ...

AIR CYCLE MACHINE COMPRESSOR HOUSING

An air cycle machine (ACM) compressor housing includes a body including an exterior surface and an interior portion. An inlet is integrally formed with the body. The inlet includes an inlet passage fluidically connected with the interior portion. An outlet is integrally formed with the body. The outlet includes an outlet passage fluidically connected with the interior portion. A bypass is integrally formed with the body. The bypass includes a bypass passage fluidically connected with the interior portion. The bypass passage includes a first end portion extending outwardly from the body, a second end portion and an intermediate portion extending therebetween. The first end portion includes a wall thickness of between about 0.330-inch (8.382-mm) and 0.390-inch (9.906-mm). 1. An air cycle machine (ACM) compressor housing comprising:a body including an exterior surface and an interior portion;an inlet integrally formed with the body, the inlet including an inlet passage fluidically connected with the interior portion;an outlet integrally formed with the body, the outlet including an outlet passage fluidically connected with the interior portion; anda bypass integrally formed with the body, the bypass including a bypass passage fluidically connected with the interior portion, the bypass passage including a first end portion extending outwardly from the body, a second end portion and an intermediate portion extending therebetween, the first end portion including a wall thickness of between about 0.330-inch (8.382-mm) and 0.390-inch (9.906-mm).2. The ACM compressor housing according to claim 1 , wherein the first end portion includes a wall thickness of about 0.360-inch (9.144-mm).3. The ACM compressor housing according to claim 1 , further comprising: a fillet extending from the body to the first end portion of the bypass claim 1 , the fillet including a radius of between about 0.970-inches (25.638-mm) and 1.030 inches (26.162-mm).4. The ACM compressor housing according ...

BEARING STRUCTURE

A bearing structure includes: a rotation member including a plurality of extended portions extending radially outward from a shaft portion and arranged separated away from each other in an axial direction of the shaft portion; and a bearing member in which a counterface surface facing one of the plurality of extended portions in the axial direction is included in one or a plurality of main bodies. 1. A bearing structure comprising:a rotation member including a shaft portion and a plurality of extended portions extending radially outward from the shaft portion and arranged separated away from each other in an axial direction of the shaft portion;a large diameter portion having a diameter larger than a diameter of the shaft portion and including the extended portion on each of one side and another side in the axial direction, the large diameter portion being included in the rotation member;a bearing member including a plurality of main bodies through which the shaft portion is inserted, the plurality of main bodies being arranged on each of one side and another side in the axial direction with respect to the large diameter portion, the plurality of main bodies including a counterface surface facing one of the plurality of extended portions in the axial direction; anda turbine housing in which a through hole is formed, the through hole in which the bearing member is arranged.2. The bearing structure according to claim 1 , wherein the rotation member includes a pair of the extended portions arranged so as to face each other in the axial direction.3. The bearing structure according to claim 1 , wherein the bearing member includes the main body located between a pair of the extended portions arranged so as to face each other in the axial direction claim 1 , the main body including a pair of counterface surfaces facing the pair of extended portions claim 1 , respectively.4. The bearing structure according to claim 2 , wherein the bearing member includes the main body ...

CONCURRENT ROCKET ENGINE PRE-CONDITIONING AND TANK LOADING

Concurrent rocket engine pre-conditioning and tank filling is disclosed. A disclosed example apparatus includes an inlet valve to supply a rocket propellant tank that is associated with a rocket engine with rocket propellant, and a flow director to direct at least a portion of a flow of the rocket propellant from the inlet valve to a chill line of the rocket engine to thermally condition the rocket engine as the rocket propellant tank is being filled with the rocket propellant. 1. An apparatus comprising:an inlet valve to supply a rocket propellant tank that is associated with a rocket engine with rocket propellant; anda flow director to direct at least a portion of a flow of the rocket propellant from the inlet valve to a chill line of the rocket engine to thermally condition the rocket engine as the rocket propellant tank is being filled with the rocket propellant.2. The apparatus as defined in claim 1 , wherein the rocket engine is disposed between the rocket propellant tank and the inlet valve.3. The apparatus as defined in claim 2 , further including a propellant line disposed between an engine bleed line and a pre-valve associated with an inlet of the rocket propellant tank.4. The apparatus as defined in claim 1 , further including a bypass valve disposed between the chill line and the rocket propellant tank to drive flow toward the chill line.5. The apparatus as defined in claim 4 , further including a valve controller to vary a degree to which the bypass valve is opened based on an amount of the rocket propellant in the rocket propellant tank.6. The apparatus as defined in claim 1 , further including a bypass branch that extends from an exit of an engine bleed line to a bypass inlet valve of the rocket propellant tank.7. The apparatus as defined in claim 6 , further including a t-shaped juncture defined by the bypass branch claim 6 , wherein an overboard bleed valve is disposed at a first branch of the t-shaped juncture claim 6 , and wherein the bypass inlet ...

AIR INTAKE SLEEVE FOR AN AIRCRAFT TURBOPROP ENGINE

Air intake sleeve () for an aircraft turboprop engine, comprising an air bleed duct () that is oriented substantially along a first axis (A) and a duct () for supplying air to a compressor, which duct is oriented substantially along a second axis (B), at a distance from the first axis and substantially in parallel with the first axis, said air bleed duct and said supply duct being interconnected by an intermediate duct () having a general S shape, when viewed from the side, characterised in that said intermediate duct comprises, on at least one of the walls thereof, means for sucking air into the flow path of the intermediate duct, said suction means () being positioned and/or designed to suck a boundary layer of said flow path. 1. Air intake sleeve for an aircraft turboprop engine , comprising an air bleed duct that is oriented substantially along a first axis and a duct for supplying air to a compressor , which duct is oriented substantially along a second axis , at a distance from the first axis and substantially in parallel with the first axis , said air bleed duct and said supply duct being interconnected by an intermediate duct having a general S shape , when viewed from the side , said intermediate duct comprising , on at least one of the walls thereof , means for sucking air from the flow path of the intermediate duct , said suction means being positioned and/or designed to suck a boundary layer of said flow path , characterised in that said suction means are designed to produce the suction in a shedding zone of the boundary layer and in that said means are designed to be connected to an air conditioning circuit of a cabin of the aircraft.2. Sleeve according to claim 1 , wherein said suction means are positioned on an air-flow-diverting wall of said intermediate duct.3. Sleeve according to claim 2 , wherein said diverting wall has a tangent that is inclined relative to said first and second axes.4. Sleeve according to claim 1 , wherein it comprises an air- ...

COMPRESSOR AND GAS TURBINE

A compressor includes: a plurality of vanes at a vane stage provided to a rotor casing demarcating the primary duct; an air bleed chamber casing that demarcates an air bleed chamber interconnecting with the primary duct; and an air bleed tubing connected to the air bleed chamber casing. Of the plurality of vanes, when a plurality of vanes positioned at a region including the position in the peripheral direction corresponding to the air bleed tubing are a first vane group and a plurality of vanes other than the first vane group are a second vane group, the spacing between the ends at the outside in the radial direction of the vanes that are adjacent in the first vane group is closer than the spacing between the ends at the outside in the radial direction of the vanes that are adjacent in the second vane group. 1. A compressor comprising:a rotor which rotates around an axis;a rotor casing which surrounds the rotor from an outer peripheral side thereof to allow a main flow path of a fluid to be defined between the rotor casing and the rotor;a plurality of stator blades which are provided with intervals therebetween in a circumferential direction so as to be directed toward an inside in a radial direction from the rotor casing;a bleed chamber casing which is provided on the outer peripheral side of the rotor casing and defines a bleed chamber that communicates with the main flow path via a slot that is formed to extend in the circumferential direction on a downstream side of the stator blade; anda pipe which is connected to the bleed chamber casing from an outer peripheral side thereof and has a bleed flow path formed therein, the bleed flow path guiding the fluid in the bleed chamber to an outside,wherein, among the plurality of stator blades, when a plurality of stator blades positioned in a region including a circumferential position corresponding to the pipe are defined as a first stator blade group and a plurality of stator blades excluding the first stator blade ...

VARIABLE PRESSURE AIR SUPPLY

The present disclosure relates to engine buffer systems. An engine buffer system may include a low pressure supply line and a high pressure supply line. A continuously variable valve may be coupled to and/or in fluid communication with the low pressure supply line and the high pressure supply line. The continuously variable valve may be adjusted to supply any pressure between a pressure of the low pressure supply line and a pressure of the high pressure supply line to a buffer line. 1. A buffer system comprising:a low pressure supply line;a high pressure supply line; anda variable valve in fluid communication with the low pressure supply line and the high pressure supply line, wherein the variable valve comprises a low pressure orifice and a high pressure orifice.2. The buffer system of claim 1 , further comprising an actuator configured to translate the variable valve.3. The buffer system of claim 1 , wherein the variable valve is configured to allow pressurized air from the low pressure supply line and the high pressure supply line to pass through the variable valve simultaneously.4. The buffer system of claim 1 , wherein a distance between the low pressure orifice and the high pressure orifice is greater than a distance between the low pressure supply line and the high pressure supply line.5. The buffer system of claim 1 , wherein the low pressure supply line is in fluid communication with at least one of a low pressure compressor and a stage between the low pressure compressor and a high pressure compressor.6. The buffer system of claim 1 , wherein the high pressure supply line is in fluid communication with an intermediate stage of a high pressure compressor.7. The buffer system of claim 1 , wherein the buffer system is configured to provide pressurized air to an engine bearing compartment.8. The buffer system of claim 1 , wherein claim 1 , in response to a desired pressure being between a pressure in the low pressure supply line and a pressure in the high ...

Variable-flow-rate valve mechanism and turbocharger

In a variable-flow-rate valve mechanism, a valve is fitted into an attachment hole of an attachment member. The valve allows for play with the attachment member, and includes a valve body provided with a valve surface. A valve shaft is integrally formed in the center of a head portion of the valve body. A stopping member is provided to a leading end portion of the valve shaft. A leaf spring is provided to the valve shaft. The leaf spring includes a folded-back portion formed by bending. An insertion hole is formed in one end portion of the leaf spring. A cutout portion is formed in another end portion of the leaf spring. The one end portion of the leaf spring is fixed to the attachment member and the other end portion of the leaf spring is pressed to the head portion of the valve body.

NACELLE WITH THRUST REVERSER

A nacelle for an aircraft having a thrust reverser including a translating cowl and a blocking panel. The blocking panel is connected by a pivot connection and is connected to the translating cowl by a sliding connection including an engagement member slidingly engaged in a track, the track having a first portion parallel to the longitudinal axis of the outer casing and a second portion non-parallel to the longitudinal axis. Motion of the translating cowl along the longitudinal axis slides the engagement member within the track, and is performed without moving the blocking panel when the engagement member is within the first portion of the track. The motion of the translating cowl causes the blocking panel to pivot about the pivot connection when the engagement member is within the second portion of the track. 1. A nacelle for an aircraft , comprising:an outer casing having a longitudinal axis; a translating cowl, the outer casing and translating cowl cooperating to define an outer boundary of a bypass passage configured to surround an engine, the translating cowl movable relative to the outer casing along the longitudinal axis between a closed position and an open position, the translating cowl in the open position being axially spaced from the outer casing so as to define an opening between the translating cowl and the outer casing; and', 'a blocking panel pivotable about a pivot connection between a stowed position and a pivoted position, the blocking panel in the pivoted position extending within the bypass passage adjacent the opening to direct a flow toward the opening;', 'wherein the blocking panel is connected to the translating cowl by a sliding connection including an engagement member slidingly engaged in a track, the track having a first portion parallel to the longitudinal axis and a second portion non-parallel to the longitudinal axis;, 'a thrust reverser includingwherein motion of the translating cowl along the longitudinal axis slides the engagement ...

GAS-TURBINE SYSTEM

A gas-turbine system of a power plant, the system having: a compressor; an annular combustion chamber having a burner; and a turbine coupled to the compressor. The gas turbine system has a bypass line for compressed air, the line bypassing the burner and having a flow connection parallel thereto. 110.-. (canceled)11. A gas turbine system of a power plant , comprising:a compressor,an annular combustion chamber comprising a burner, anda turbine which is coupled to the compressor, and having a bypass line for compressed air bypassing the burner and connected flow-wise in parallel therewith,wherein the bypass line opens in an annular groove,wherein the annular groove is covered by thermal tiles.12. The gas turbine system as claimed in claim 11 ,wherein the annular groove is covered by a plate comprising a number of openings.13. The gas turbine system as claimed in claim 11 ,wherein the annular groove is introduced into the annular combustion chamber.14. The gas turbine system as claimed in claim 11 ,wherein the bypass line comprises an axial slot introduced on the hub side of the annular combustion chamber.15. The gas turbine system as claimed in claim 11 ,wherein the bypass line opens in at least two annular grooves.16. The gas turbine system as claimed in claim 15 ,wherein the annular combustion chamber comprises an outer shell and a hub-side inner shell,wherein each shell comprises one of the annular grooves.17. The gas turbine system as claimed in claim 11 ,wherein the bypass line comprises a valve coupled for signaling to a control circuit.18. A method for operating a gas turbine system as claimed in claim 17 , the method comprising:supplying the bypass line with air as a function of power demand, andsupplying the valve with a control command. This application is the US National Stage of International Application No. PCT/EP2015/055765 filed Mar. 19, 2015, and claims the benefit thereof. The International Application claims the benefit of German Application No. DE ...

HEAT CYCLE SYSTEM

A heat cycle system includes a cooling circuit and a Rankine cycle circuit in which an organic medium circulates. The Rankine cycle circuit includes an evaporator, an expander, and a condenser. Before warm-up of an engine, a control device executes a warm-up mode in which the organic medium is circulated through the condenser, the expander and the evaporator in sequence; after the warm-up of the engine, the control device executes a waste heat recovery mode in which the organic medium is circulated through the evaporator, the expander and the condenser in sequence. In the warm-up mode, by supplying energy to the expander, the control device compresses the organic medium passing through the condenser and supplies the compressed organic medium to the evaporator; in the waste heat recovery mode, by depressurizing the organic medium passing through the evaporator by the expander, the control device recovers the energy generated by the expander. 1. A heat cycle system , comprising:a cooling circuit in which cooling water performing heat exchange with an internal combustion engine and exhaust of the internal combustion engine circulates; a first heat exchanger performing heat exchange between an organic medium having a lower boiling point than the cooling water and the cooling water of the cooling circuit,', 'an expander depressurizing the organic medium passing through the first heat exchanger and generating energy, and', 'a second heat exchanger performing heat exchange between the organic medium and outside air; and, 'a Rankine cycle circuit, comprisinga control device operating the Rankine cycle circuit in a warm-up mode in which the organic medium is circulated through the second heat exchanger, the expander, and the first heat exchanger in this order before warm-up of the internal combustion engine, and operating the Rankine cycle circuit in a waste heat recovery mode in which the organic medium is circulated through the first heat exchanger, the expander, and the ...

GAS TURBINE

A gas turbine includes a compressor; a combustor; a turbine configured to drive a rotational shaft of the compressor using combustion gas generated by the combustor; a cooling device configured to generate cooling air by bleeding compressed air from the compressor and cooling the compressed air, and to supply the cooling air to the turbine along the rotational shaft; a pressurizing device configured to increase pressure of the cooling air; a pressurizing device diffuser configured to provide a passage continuing in a turbine circumferential direction, on the outer side in the turbine radial direction to guide the cooling air having the increased pressure to the outer side of the pressurizing device; and a manifold disposed between the pressurizing device diffuser and a plurality of turbine vanes so that a ring-shaped passage communicates with the passage in the pressurizing device diffuser and a cooling passage provided inside each turbine vane. 1. A gas turbine comprising:a compressor configured to rotate about a rotational shaft to generate compressed air;a combustor configured to generate combustion gas using the compressed air generated by the compressor;a turbine configured to drive the rotational shaft in rotation using the combustion gas generated by the combustor;a cooling device configured to generate cooling air by bleeding the compressed air from the compressor and cooling the compressed air, and to supply the cooling air to the turbine along the rotational shaft;a pressurizing device disposed between the cooling device and the turbine to increase pressure of the cooling air toward an outer side in a turbine radial direction as the rotational shaft rotates;a pressurizing device diffuser configured to provide a passage continuing in a turbine circumferential direction, on the outer side of the pressurizing device in the turbine radial direction so as to guide the cooling air having pressure increased by the pressurizing device to the outer side of the ...

VARIABLE FAN NOZZLE USING SHAPE MEMORY MATERIAL

A gas turbine engine includes a fan, a nacelle arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow from the fan. A nozzle associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position. 1. A gas turbine engine comprising:a fan;a nacelle arranged about the fan;an engine core at least partially within the nacelle, the engine core having a compressor and a turbine;a fan bypass passage downstream of the fan between the nacelle and the gas turbine engine, for conveying a bypass airflow from the fan;a nozzle section for controlling the bypass airflow, wherein the nozzle section includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position.2. The gas turbine engine recited in claim 1 , wherein the shape memory material comprises a threshold temperature that corresponds to a reversible change between the first solid state phase and the second solid state phase.3. The gas turbine engine recited in claim 1 , wherein the shape memory material comprises a nickel-titanium alloy.4. The gas turbine engine recited in claim 1 , wherein the shape memory material comprises a material selected from a copper alloy claim 1 , a nickel alloy claim 1 , a cobalt alloy claim 1 , a manganese alloy claim 1 , a copper-aluminum alloy claim 1 , a copper-zinc-aluminum alloy claim 1 , and combinations thereof.5. The gas turbine engine recited in claim 1 , wherein the nozzle includes tabs that include the shape memory material.6. The gas turbine engine recited in claim 5 , wherein the tabs extend from the nacelle in a ...

CONTROLLING A TURBOCHARGER SYSTEM

A turbocharger control system includes a turbine; a fluid source of a pressurized fluid; an input valve fluidly coupled between the fluid source and an input of the turbine; a bypass valve fluidly coupled between the fluid source and an output of the turbine; a rotating machine operatively coupled to the turbine and configured to move a working fluid; and a control system communicably coupled to the input valve and the bypass valve. The control system is configured to perform operations including determining a level of the pressurized fluid in the fluid source; determining at least one of a flow rate or a pressure of a working fluid moved by the rotating machine; and operating the input valve and the bypass valve to change an operating state of the turbine from a first operating state to a second operating state. 1. A turbocharger control system , comprising:a turbine;a fluid source fluidly coupled to the turbine, the fluid source comprising a pressurized fluid;at least one input valve fluidly coupled between the fluid source and an input of the turbine;at least one bypass valve fluidly coupled between the fluid source and an output of the turbine;a rotating machine operatively coupled to the turbine and configured to move a working fluid; and determining a level of the pressurized fluid in the fluid source;', 'determining at least one of a flow rate or a pressure of a working fluid moved by the rotating machine; and', 'based on at least one of the determined fluid source level and the at least one of the determined flow rate or pressure, operating the at least one input valve and the at least one bypass valve to change an operating state of the turbine from a first operating state to a second operating state., 'a control system communicably coupled to the at least one input valve and the at least one bypass valve, the control system configured to perform operations comprising2. The turbocharger control system of claim 1 , wherein the pressurized fluid comprises ...

DIRT MITIGATION IN A GAS TURBINE ENGINE

An aspect includes a dirt mitigation system for a gas turbine engine. The dirt mitigation system includes a plurality of bleeds of the gas turbine engine and a control system configured to determine a particulate ingestion estimate indicative of dirt ingested in the gas turbine engine. The control system is further configured to determine one or more operating parameters of the gas turbine engine and alter a bleed control schedule of the gas turbine engine to purge at least a portion of the dirt ingested in the gas turbine engine through one or more of the bleeds of the gas turbine engine based on the particulate ingestion estimate and the one or more operating parameters. 1. A dirt mitigation system for a gas turbine engine , the dirt mitigation system comprising:a plurality of bleeds of the gas turbine engine; and determine a particulate ingestion estimate indicative of dirt ingested in the gas turbine engine;', 'determine one or more operating parameters of the gas turbine engine; and', 'alter a bleed control schedule of the gas turbine engine to purge at least a portion of the dirt ingested in the gas turbine engine through one or more of the bleeds of the gas turbine engine based on the particulate ingestion estimate and the one or more operating parameters., 'a control system configured to2. The dirt mitigation system of claim 1 , wherein the particulate ingestion estimate is determined based on a plurality of particle sensor data from the gas turbine engine.3. The dirt mitigation system of claim 2 , wherein the particle sensor data is received from one or more particle sensors at an inlet of a low pressure compressor of the gas turbine engine downstream of a fan of the gas turbine engine.4. The dirt mitigation system of claim 2 , wherein the control system is further configured to determine whether the particle sensor data indicates that a dirt concentration is above a threshold claim 2 , an engine deterioration is within a deterioration limit claim 2 , and ...

ADAPTIVE BLEED SCHEDULE IN A GAS TURBINE ENGINE

An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions. 1. A system for a gas turbine engine , the system comprising:one or more bleeds of the gas turbine engine; and check one or more activation conditions of a dirt rejection mode in the gas turbine engine;', 'adjust a bleed control schedule of the gas turbine engine to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions;', 'check one or more deactivation conditions of the dirt rejection mode in the gas turbine engine; and', 'deactivate the dirt rejection mode to fully close the one or more bleeds based on the one or more deactivation conditions., 'a control system configured to2. The system of claim 1 , wherein the dirt rejection mode is activated based on meeting all of the one or more activation conditions claim 1 , and the dirt rejection mode is deactivated based on meeting at least one of the one or more deactivation conditions.3. The system of claim 1 , wherein the one or more activation conditions comprise detecting that the gas turbine engine is incorporated in an aircraft on the ground and an engine speed command of the gas turbine engine is less than a full maximum takeoff thrust setting with a thrust margin.4. The system of claim 3 , wherein the ...

AIRCRAFT HAVING AN ENGINE AND A COOLING SYSTEM

An aircraft having an engine, a tank, devices to be heated, air intakes, a heat exchanger, a pipe connecting the air intakes and the devices to be heated by passing through the heat exchanger, a fuel pipe connected between the tank and the combustion chamber of the engine, and an air pipe which feeds the heat exchanger from the fan duct. The aircraft has an additional heat exchanger installed on the pipe between the heat exchanger and the air intakes, and a diversion pipe that passes through the additional heat exchanger and is connected at each end to the fuel pipe. The use of the fuel to cool the air makes it possible to use a smaller and therefore less bulky heat exchanger. 1. An aircraft comprising:an engine having a high-pressure compressor with multiple compression stages, a combustion chamber and a fan duct,a fuel tank containing fuel,devices to be heated,a first air intake configured to draw, from the high-pressure compressor, air at a low pressure or at an intermediate pressure, a second air intake configured to draw, from the high-pressure compressor, air at a high pressure,a first heat exchanger,a first pipe which passes through the first heat exchanger and feeds the devices to be heated downstream of the first heat exchanger, wherein, upstream of the first heat exchanger, the first pipe is divided into two sub-pipes, one of which is fluidically connected to the first air intake and the other of which is fluidically connected to the second air intake,a fuel pipe fluidically connected between the fuel tank and the combustion chamber of the engine, anda first air pipe which feeds the first heat exchanger with air drawn from the fan duct,an additional heat exchanger installed on the first pipe upstream of the first heat exchanger and downstream of the first and second air intakes,a diversion pipe, a first branch of which is fluidically connected between the fuel pipe and an inlet of the additional heat exchanger and a second branch of which is fluidically ...

COOLING OF GAS TURBINE ENGINE ACCESSORIES

A gas turbine engine for an aircraft is provided. The engine includes an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The engine further includes core casings surrounding the engine core. The engine further includes an aerodynamic cowl which surrounds the core casings. The engine further includes a propulsive fan located upstream of the engine core, the fan generating a core airflow which enters the core engine and a bypass airflow which enters a bypass duct surrounding the aerodynamic cowl. The engine further includes one or more engine accessories mounted in a space between the core casings and the aerodynamic cowl. 1. A gas turbine engine for an aircraft , the engine including:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;core casings surrounding the engine core;an aerodynamic owl which surrounds the core casings;a propulsive fan located upstream of the engine core, the fan generating a core airflow which enters the core engine and a bypass airflow which enters a bypass duct surrounding the aerodynamic cowl; andone or more engine accessories mounted in a space between the core casings and the aerodynamic cowl;wherein the gas turbine engine further includes:one or more ventilation inlets which receive a portion of the bypass airflow as a cooling flow for the engine accessories, the ventilation inlets being configured to convert at least a portion of the kinetic energy of the bypass flow received therein into a pressure rise such that the pressure of the cooling flow is increased relative to the pressure of the bypass airflow; anda manifold in fluid communication with the ventilation inlets to collect the cooling flow therefrom, the manifold having plural exhaust holes therefrom through which the cooling flow leaves the manifold to impinge upon and thereby cool the engine accessories.2. A gas turbine engine according to claim 1 , wherein the ...

REDUCED NOISE TURBOFAN AIRCRAFT ENGINE

The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. For the first stage the mean radius r of a stator vane expressed in inch divided by the number of stator vanes is at least 0.18. 2. The turbofan aircraft engine of claim 1 , wherein the bypass ratio is at least 8.3. The turbofan aircraft engine of claim 1 , wherein the bypass ratio is at least 9.4. The turbofan aircraft engine of claim 1 , wherein r/n of the first stage is at least 0.19.5. The turbofan aircraft engine of claim 1 , wherein r/n of the first stage is at least 0.195.6. The turbofan aircraft engine of claim 1 , wherein en of the first stage is at least 0.20.7. The turbofan aircraft engine of claim 3 , wherein r/n of the first stage is at least 0.20.8. The turbofan aircraft engine of claim 1 , wherein r/n of the second stage is at least 0.17.9. The turbofan aircraft engine of claim 1 , wherein r/n of the second stage is at least 0.175.10. The turbofan aircraft engine of claim 1 , wherein r/n of the second stage is at least 0.18.11. The turbofan aircraft engine of claim 9 , wherein r/n of the first stage is at least 0.19.12. The turbofan aircraft engine of claim 1 , wherein the second turbine comprises not more than three stages.13. The turbofan aircraft engine of claim 1 , wherein r/n of the first stage is not higher than 0.26.14. The turbofan aircraft engine of claim 1 , wherein n of the first stage ranges from 45 to 80.15. The turbofan aircraft engine of claim 1 , wherein n of the first ...

INTERCOOLED COOLING AIR TAPPED FROM PLURAL LOCATIONS

A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed. 1. A gas turbine engine comprising;a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations;a turbine section having a high pressure turbine;a first tap tapping air from at least one of said more upstream locations in said main compressor section, passing said tapped air through a heat exchanger and then to a cooling compressor, said cooling compressor compressing air downstream of said heat exchanger; anda second tap tapping air from a location closer to said downstream most end than the location(s) of said first tap, and said first and second tap mixing together and being delivered into said high pressure turbine.2. The gas turbine engine as set forth in claim 1 , wherein a main fan delivers bypass air into a bypass duct and into said main compressor section and said heat exchanger positioned within said bypass duct to be cooled by bypass air.3. The gas turbine engine as set forth in claim 1 , wherein air temperatures at said downstream most location of said high pressure compressor are greater than or equal to 1350° F.4. The gas turbine engine as set forth in claim 1 , wherein the second tap is at said downstream most end.5. The gas turbine engine as set forth in claim 4 , wherein air from said first ...

EXHAUST-GAS TURBOCHARGER

An exhaust-gas turbocharger with a compressor with a charge pressure control device. An actuator is provided which moves a control rod arrangement. The control rod arrangement is in turn rotatably connected to a lever. The charge pressure control device is actuated by the lever. A pin is arranged on the lever. The control rod arrangement has a bushing which is pushed onto the pin. The connection between pin and bushing constitutes a rotatable connection for enabling the lever to be actuated by the control rod arrangement. This arrangement is very simple to produce and very simple to assemble because the bushing must merely be pushed onto the pin. 11. An exhaust-gas turbocharger () comprising:{'b': 5', '7, 'a compressor () with a compressor wheel (),'}{'b': 2', '4, 'a turbine () with a turbine wheel (),'}{'b': '9', 'a charge pressure control device (),'}{'b': 13', '9, 'an actuator () for actuating the charge pressure control device (),'}{'b': 12', '13, 'a control rod arrangement () that can be moved by the actuator (), and'}{'b': 11', '9', '9, 'a lever (), which is connected to the charge pressure control device (), for opening and closing the charge pressure control device (),'}{'b': 12', '11, 'the control rod arrangement () being pivotably connected to the lever (),'}{'b': 12', '16', '11', '18', '18', '16, 'the control rod arrangement () having a bushing () and the lever () having a pin (), the pin () being pivotably inserted in the bushing (), and'}{'b': 12', '14', '15', '16', '15', '16', '15, 'the control rod arrangement () having a rod (), on the end of which there is arranged a guide piece (), the bushing () being a constituent part of the guide piece (), or the bushing () being directly and fixedly connected to the guide piece (),'}{'b': 15', '20', '14', '20', '15, 'wherein the guide piece () has a partially trough-shaped cutout (), the rod () being placed into the cutout () and welded to, pressed together with or brazed to the guide piece ().'}218. The ...

FUEL CONTROL SYSTEM

A bypass valve for a fuel control includes a sleeve, a spool, and a biasing member. The sleeve has a body extending between a first end and a second. The first end defines a first bypass valve port. The body defines a second bypass valve port, a third bypass valve port, and a shutoff port. The second end defines a fourth bypass valve port. The spool is received within the inner bore and is movable between a first position and a second position. The biasing member biases the spool toward the first position. 1. A fuel control system comprising: a first shutoff valve port fluidly connected to a metering valve outlet,', 'a second shutoff valve port fluidly connected to an engine, and', 'a third shutoff valve port fluidly connected to a damping orifice; and, 'a shutoff valve having'} a first bypass valve port fluidly connected to the metering valve outlet,', 'a second bypass valve port fluidly connected to a fuel pump inlet,', 'a third bypass valve port fluidly connected to a fuel pump outlet,', 'a fourth bypass valve port selectively fluidly connected to the fuel pump inlet, the fuel pump outlet, and the third shutoff valve port, and', 'a shutoff port fluidly connected to the third shutoff valve port, in response to an actuator fluidly connecting the fourth bypass valve port to the fuel pump inlet, the bypass valve moves from a closed position toward an open position., 'a bypass valve having'}2. The fuel control system of claim 1 , wherein in response to the bypass valve moving from the closed position toward the open position a fuel pressure supplied to the engine decreases.3. The fuel control system of claim 1 , wherein a shutoff port diameter is greater than a damping orifice diameter.4. The fuel control system of claim 1 , wherein the bypass valve is held in the closed position by at least one of a fuel pump outlet supply pressure and a bypass valve biasing member.5. The fuel control system of claim 1 , wherein the bypass valve is moved from the closed position ...

INLET BLEED HEAT CONTROL SYSTEM

The present application provides an inlet bleed heat control system for a compressor of a gas turbine engine. The inlet bleed heat control system provides an inlet bleed heat manifold and an ejector in communication with the inlet bleed heat manifold such that the ejector is in communication with a flow of compressor discharge air and a flow of ambient air. 1. An inlet bleed heat control system for a compressor of a gas turbine engine , comprising:an inlet bleed heat manifold; andan ejector in communication with the inlet bleed heat manifold;the ejector in communication with a flow of compressor discharge air and a flow of ambient air.2. The inlet bleed heat control system of claim 1 , wherein the ejector is in communication with a filter house line with the flow of ambient air.3. The inlet bleed heat control system of claim 2 , wherein the filter house line comprises a filter house line valve thereon.4. The inlet bleed heat control system of claim 1 , wherein the ejector is in communication with a compressor discharge line with the flow of compressor discharge air.5. The inlet bleed heat control system of claim 4 , wherein the compressor discharge line comprises a compressor discharge valve thereon.6. The inlet bleed heat control system of claim 5 , wherein the compressor discharge line comprises an ejector valve thereon.7. The inlet bleed heat control system of claim 1 , further comprising an ejector bypass line positioned between the compressor and the inlet bleed heat manifold.8. The inlet bleed heat control system of claim 7 , wherein the bypass line comprises a bypass valve thereon.9. The inlet bleed heat control system of claim 1 , wherein the inlet bleed heat manifold comprises a plurality of acoustic nozzles.10. The inlet bleed heat control system of claim 1 , further comprising a controller.11. The inlet bleed heat control system of claim 10 , wherein the controller comprises anti-icing signals.12. The inlet bleed heat control system of claim 10 , wherein ...

Fluid recirculation turbine system

A rotor rotatably mounted within a turbocharger housing includes a turbine wheel and a shaft. The shaft connects the turbine wheel. The hub defines a turbine-wheel back-disk surface facing the portion of the housing containing the bearings, and the hub defines a blade-side surface. The turbine hub and the housing define a turbine-wheel back-disk cavity. The turbine hub forms a ring-shaped primary axial protrusion extending circularly around the turbine-wheel back-disk surface into a circular channel in the housing. The circular channel leads into a bypass that bypasses the turbine blades. A relief flow valve is placed in the bypass. The relief control valve is controlled to open when the bypass pressure is above a cutoff pressure, and close when it is below the cutoff pressure.

ADAPTIVE BLEED SCHEDULE IN A GAS TURBINE ENGINE

An aspect includes a system for a gas turbine engine. The system includes one or more bleeds of the gas turbine engine and a control system configured to check one or more activation conditions of a dirt rejection mode in the gas turbine engine. A bleed control schedule of the gas turbine engine is adjusted to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions. One or more deactivation conditions of the dirt rejection mode in the gas turbine engine are checked. The dirt rejection mode is deactivated to fully close the one or more bleeds based on the one or more deactivation conditions. 1. A system for a gas turbine engine , the system comprising:one or more bleeds of the gas turbine engine; and check one or more activation conditions of a dirt rejection mode in the gas turbine engine;', 'read a default bleed control schedule from the memory system as a bleed control schedule;', 'determine that both a dirt concentration is above a threshold and the gas turbine engine is within an operability limit;', 'in response to determining that both the dirt concentration is above a threshold and the gas turbine engine is within the operability limit, adjust the bleed control schedule of the gas turbine engine to extend a time to hold the one or more bleeds of the gas turbine engine partially open at a power setting above a threshold based on the one or more activation conditions;', 'check one or more deactivation conditions of the dirt rejection mode in the gas turbine engine; and', 'deactivate the dirt rejection mode to fully close the one or more bleeds based on the one or more deactivation conditions and restore the bleed control schedule of the gas turbine engine by updating the bleed control schedule with the default bleed control schedule from the memory system., 'a control system comprising processing circuitry and a memory system, the memory system ...

Thermal Gradient Attenuation Structure to Mitigate Rotor Bow in Turbine Engine

Embodiments are generally provided of a gas turbine engine including a rotor assembly comprising a shaft extended along a longitudinal direction, in which a compressor rotor and a turbine rotor are each coupled to the shaft; a casing surrounding the rotor assembly, in which the casing defines a first opening radially outward of the compressor rotor, the turbine rotor, or both, and a second opening radially outward of the compressor rotor, the turbine rotor, or both; a first manifold assembly coupled to the casing at the first opening; a second manifold assembly coupled to the casing at the second opening, in which the first manifold, the casing, and the second manifold together define a thermal circuit in thermal communication with the rotor assembly; and a fluid flow device in fluid communication with the first manifold assembly, in which the fluid flow device provides a flow of fluid to the first manifold assembly and through the thermal circuit, and further wherein the flow of fluid egresses the thermal circuit at the second manifold assembly. 1. A gas turbine engine defining an axial centerline and a longitudinal direction extended co-directional thereto and a radial direction extended from the axial centerline , the engine comprising:a rotor assembly comprising a shaft extended along a longitudinal direction, wherein a compressor rotor and a turbine rotor are each coupled to the shaft;a casing surrounding the rotor assembly, wherein the casing defines a first opening radially outward of the compressor rotor, the turbine rotor, or both, and a second opening radially outward of the compressor rotor, the turbine rotor, or both;a first manifold assembly coupled to the casing at the first opening;a second manifold assembly coupled to the casing at the second opening, wherein the first manifold, the casing, and the second manifold together define a thermal circuit in thermal communication with the rotor assembly; anda fluid flow device in fluid communication with the ...

ASSISTED ENGINE START BLEED SYSTEM

A system for bleeding air from a core flow path of a gas turbine engine includes a bleed valve in a bleed air duct configured to receive bleed air from a first entrance point to the core flow path into the bleed air duct; a pressurized air valve in a pressurized air duct configured to receive pressurized air from a second entrance point to the core flow path, the pressurized air at a pressure greater than that received into the first entrance point; an eductor outlet from the pressurized air duct located in the bleed air duct; and a control system operable to control operation of the bleed valve and the pressurized air valve. 1. A system for bleeding air from a core flow path of a gas turbine engine , comprising:a bleed valve in a bleed air duct configured to receive bleed air from a first entrance point to the core flow path into the bleed air duct;a pressurized air valve in a pressurized air duct configured to receive pressurized air from a second entrance point to the core flow path, the pressurized air at a pressure greater than that received into the first entrance point;an eductor outlet from the pressurized air duct located in the bleed air duct; anda control system operable to control operation of the bleed valve and the pressurized air valve.2. The system as recited in claim 1 , wherein the bleed air duct is of a larger diameter than the pressurized air duct.3. The system as recited in claim 1 , wherein the controller is a FADEC.4. The system as recited in claim 1 , wherein the first entrance point is positioned proximate a compressor section of the gas turbine engine.5. The system as recited in claim 1 , wherein the first entrance point is positioned proximate a low pressure compressor section of the gas turbine engine.6. The system as recited in claim 1 , wherein the first entrance point is positioned upstream of a low pressure compressor section of the gas turbine engine.7. The system as recited in claim 1 , wherein the second entrance point is ...

TURBINE INCLUDING FLUE GAS RECIRCULATION COMBUSTOR

A flue gas recirculation combustor includes a combustor chamber configured such that fuel and combustion gas are injected therein to cause combustion and having a nozzle-side end and a combustor outlet, a nozzle can connected to the nozzle-side end of the combustor chamber, a plurality of nozzles disposed in the nozzle can and configured such that an injection direction thereof is directed to a side of the combustor chamber, and a sleeve disposed in a premixing space defined between the nozzle can and the nozzle-side end of the combustor chamber, the sleeve including a recirculation pathway to recirculate combustion air from the combustor chamber to the premixing space. 1. A flue gas recirculation combustor comprising:a combustor chamber configured to receive fuel and combustion gas to cause combustion, the combustor chamber including a nozzle-side end and a combustor outlet;a nozzle can connected to the nozzle-side end of the combustor chamber;a plurality of nozzles disposed in the nozzle can and arranged such that an injection direction thereof is directed to a side of the combustor chamber; anda sleeve disposed in a premixing space defined between the nozzle can and the nozzle-side end of the combustor chamber, the sleeve including a recirculation pathway to recirculate combustion air from the combustor chamber to the premixing space.2. The flue gas recirculation combustor of claim 1 , wherein the recirculation pathway includes at least one of a bypass path and an ejector nozzle.3. The flue gas recirculation combustor of claim 1 , wherein the plurality of nozzles is configured to premix and inject combustion air and fuel.4. The flue gas recirculation combustor of claim 1 , wherein the recirculation pathway is provided on an inner surface of the sleeve along a circumferential direction.5. The flue gas recirculation combustor of claim 2 , wherein the recirculation pathway includes both the bypass path and the ejector nozzle claim 2 , wherein the ejector nozzle is ...

TRANSLATING CASCADE THRUST REVERSER WITH CONTROL OF BLOCKER DOOR

Aspects of the disclosure are directed to a system associated with a thrust reverser of an aircraft, comprising: a set of cascades incorporating a track, the track including a first track end and a second track end, a blocker door, and a link including a first link end coupled to the track and a second link end coupled to the blocker door, where the cascades are configured to translate between a stowed position and a deployed position to cause the first link end to traverse the track. 1. A system associated with a thrust reverser of an aircraft , comprising:a set of cascades incorporating a track, the track including a first track end and a second track end;a blocker door; anda link including a first link end coupled to the track and a second link end coupled to the blocker door,wherein the cascades are configured to translate between a stowed position and a deployed position to cause the first link end to traverse the track.2. The system of claim 1 , further comprising:a sleeve.3. The system of claim 2 , wherein the sleeve and the cascades are configured to translate via a common actuator or actuation mechanism.4. The system of claim 2 , wherein the blocker door is configured to be contained within a cavity of the sleeve when the thrust reverser is stowed.5. The system of claim 1 , wherein the first link end is configured to be located within the first track end when the thrust reverser is stowed.6. The system of claim 5 , wherein the first link end is configured to be located within the second track end when the thrust reverser is at least partially deployed in an amount greater than a threshold.7. The system of claim 6 , wherein the blocker door is configured to be in a stowed position relative to a bypass duct when the first link end is not located within the second track end.8. The system of claim 6 , wherein the blocker door is configured to be in an at least partially deployed position relative to a bypass duct when the first link end is located within the ...

AIR OIL COOLER AIRFLOW AUGMENTATION SYSTEM

An oil supply system for a gas turbine engine has a lubricant pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear system of an associated gas turbine engine. A method and apparatus are disclosed. The heat exchangers include at least an air/oil cooler wherein air is pulled across the air/oil cooler to cool oil. The air/oil cooler is provided with an ejector tapping compressed air from a compressor section to increase airflow across the air/oil cooler. 1. A lubricant supply system for a gas turbine engine comprising:a lubricant pump delivering lubricant to an outlet line that splits into at least a hot line and a cool line, said hot line being directed primarily to locations associated with the gas turbine engine that are not intended to receive cooler lubricant, and said cool line being directed to at least one component with an operating temperature lower than the locations associated with the gas turbine engine that are not intended to receive cooler lubricant; andan air/oil cooler that pulls air across said air/oil cooler to cool lubricant, wherein at least a portion of the lubricant in said cool line bypasses the air/oil cooler and is passed through a first fuel/oil cooler configured to cool the lubricant using fuel leading to a combustion section of the gas turbine engine.2. The system as recited in claim 1 , including an ejector within the air/oil cooler for augmenting air flow across the air/oil cooler selectively responsive to an operating condition not providing air flow sufficient for cooling lubricant to desired temperature.3. The system as recited in claim 2 , including selectively actuating a fan to provide airflow through to ...

WASTE GATE VALVE FOR TURBOCHARGER

A waste gate valve for a turbocharger is provided that improves control responsivity, wear resistance and the flow uniformity of diverted exhaust gas by improving the operational structure thereof. The waste gate valve includes a turbine housing in which a portion of a bypass passage is formed along the inner circumferential surface of a discharge passage. A mixer ring includes bypass apertures therein that are arranged along the bypass passage formed in the inner circumferential surface of the discharge passage. Additionally, a control valve is configured to open or close the bypass apertures based on movement thereof. 1. A waste gate valve for a turbocharger , comprising:a turbine housing that accommodates a turbine wheel therein and includes a bypass passage formed therein to divert exhaust gas introduced into the turbine wheel to a discharge passage, a portion of the bypass passage being formed along an inner circumferential surface of the discharge passage;a mixer ring formed to block a region between the bypass passage and the discharge passage and including therein bypass apertures arranged along the bypass passage formed in an inner circumferential surface of the discharge passage; anda control valve configured to switch a state of the bypass apertures to an open state or a closed state via movement thereof to regulate an opening degree of the bypass apertures.2. The waste gate valve according to claim 1 , wherein the mixer ring is fixedly inserted into a discharge pipe claim 1 , and the control valve is movably inserted between an outer circumferential surface of the mixer ring and an inner circumferential surface of the discharge pipe.3. The waste gate valve according to claim 1 , wherein the bypass passage includes:a first passage formed between an introduction pipe through which exhaust gas is introduced into the turbine housing and a discharge pipe through which exhaust gas is discharged; anda second passage that communicates with the first passage and ...

JET ENGINE COMPRISING A DRAWING SYSTEM INTENDED TO DRAW AIR FROM SAID JET ENGINE

A two-flow jet engine comprising a high-pressure compressor and a drawing system configured to draw air from said jet engine and send it to an air system, and which comprises a pressurized air supply which draws pressurized air from the high-pressure compressor, a pipe which draws outside air, a turbine of which an inlet is connected to the outlet of the pressurized air supply, and a compressor of which an inlet is connected to the pipe, and an outlet is connected to the air system, and where the rotation of the turbine drives the rotation of the compressor. This particular arrangement makes possible an energy gain by virtue of the combined action of the turbine and of the compressor. 1. A two-flow jet engine comprising a high-pressure compressor , a secondary air duct and a drawing system configured to draw air from said jet engine and send the drawn air to an air system , and which comprises:a pressurized air supply configured to draw pressurized air from the high-pressure compressor,a pipe configured to draw outside air from the secondary air duct,a turbine having a rotary shaft and of which an inlet is fluidically connected to an outlet of the pressurized air supply, anda compressor having a rotary shaft and of which an inlet is fluidically connected to the pipe, and an outlet is fluidically connected to the air system,wherein a rotation of the turbine drives the rotation of the compressor,wherein the drawing system comprises a heat exchanger,wherein the outlet of the compressor is fluidically connected to a first inlet of the heat exchanger,wherein a first outlet of the heat exchanger corresponding to the first inlet is fluidically connected to the air system,wherein an outlet of the turbine is fluidically connected to a second inlet of the heat exchanger, andwherein a second outlet of the heat exchanger corresponding to the second inlet evacuates cooled air to the outside.2. The jet engine according to claim 1 , wherein the rotary shaft of the compressor and ...

FLUID DRAIN SYSTEM FOR AN AIRCRAFT PROPULSION SYSTEM

An assembly is provided for an aircraft propulsion system. This assembly includes a first drain tube, a second drain tube, a container and a gas tube. The container fluidly couples the first drain tube to the second drain tube. The container is configured to receive fluid from the first drain tube. The gas tube is fluidly coupled with the container. The gas tube is configured to direct gas into the container for propelling the fluid received within the container into the second drain tube. 1. An assembly for an aircraft propulsion system , comprising:a first drain tube;a second drain tube;a container between and fluidly coupling the first drain tube and the second drain tube, the container configured to receive fluid from the first drain tube; anda gas tube fluidly coupled with the container, the gas tube configured to direct gas into the container for propelling the fluid received within the container into the second drain tube.2. The assembly of claim 1 , wherein the first drain tube comprises a gas turbine engine drain tube configured to receive the fluid from a component within the aircraft propulsion system.3. The assembly of claim 1 , wherein the second drain tube comprises an overboard drain tube configured to direct the fluid out of the aircraft propulsion system.4. The assembly of claim 1 , further comprising a gas source fluidly coupled with the gas tube claim 1 , the gas source configured to direct the gas through the gas tube and into the container.5. The assembly of claim 4 , wherein the gas source comprises an inlet configured to receive the gas from a bypass duct of the aircraft propulsion system.6. The assembly of claim 1 , whereinthe container is configured with an internal cavity;a first outlet portion of the first drain tube projects into the internal cavity; anda second outlet portion of the gas tube projects into the internal cavity.7. The assembly of claim 1 , wherein the container is configured to provide a visual line of sight from outside of ...

TURBINE ENGINE WITH GUIDE VANES FORWARD OF ITS FAN BLADES

A turbine engine such as a pusher fan engine is provided. This turbine engine includes a nacelle with a bypass flowpath. A fan rotor is configured to propel air out of the bypass flowpath. A plurality of guide vanes are configured to direct the air to the fan rotor. 1. A pusher fan engine , comprising:a nacelle having a bypass inlet and a nozzle, with a bypass flowpath extending from the bypass inlet to the nozzle;a pusher fan rotor including a plurality of fan blades within the bypass flowpath; anda plurality of guide vanes within the bypass flowpath and between the inlet and the fan blades.2. The pusher fan engine of claim 1 , wherein the guide vanes are pitched in an opposite direction from the fan blades.3. The pusher fan engine of claim 1 , wherein the guide vanes are pitched in a substantially equal but opposite direction from the fan blades.4. The pusher fan engine of claim 1 , wherein at least one of the fan blades is configured as a variable pitch fan blade.5. The pusher fan engine of claim 1 , wherein at least one of the guide vanes is a structural guide vane.6. The pusher fan engine of claim 1 , wherein the nacelle includes an inner casing and an outer casing claim 1 , and the inner and the outer casings are structurally tied together through a mount system which includes at least one of the guide vanes.7. The pusher fan engine of claim 1 , wherein the nacelle includes an inner casing and an outer casing claim 1 , and wherein the outer casing is structurally tied to the inner casing through at least one of the guide vanes.8. The pusher fan engine of claim 1 , further comprising a bifurcation extending through the bypass flowpath and between the inlet and the fan rotor.9. The pusher fan engine of claim 8 , wherein the bifurcation is configured as a substantially non-structural component of the pusher fan engine.10. The pusher fan engine of claim 1 , further comprising a turbine rotor and a gear train connecting the turbine rotor with the fan rotor.11. The ...

GEARED TURBOFAN ARCHITECTURE FOR REGIONAL JET AIRCRAFT

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan situated at an inlet of a bypass passage, and a core engine configured to drive the fan. The core engine includes a low pressure compressor section driven by a low pressure turbine section, and a high pressure compressor section driven by a high pressure turbine section. The fan has a fan diameter, Dfan, and the high pressure compressor section has a compressor diameter, Dcomp. The fan diameter Dfan and the compressor diameter Dcomp have an interdependence represented by a scalable ratio Dfan/Dcomp that is greater than about 4.5. 1. A gas turbine engine comprising:a fan situated at an inlet of a bypass passage; a low pressure compressor section driven by a low pressure turbine section, and', 'a high pressure compressor section driven by a high pressure turbine section; and, 'a core engine configured to drive the fan, the core engine includingwherein the gas turbine engine has a bypass ratio greater than about 10, the fan has a fan diameter, Dfan, the high pressure compressor section has a compressor diameter, Dcomp, and the fan diameter Dfan and the compressor diameter Dcomp have an interdependence represented by a scalable ratio Dfan/Dcomp that is greater than about 4.5.2. The gas turbine engine as recited in claim 1 , wherein the fan has fewer than 26 fan blades.3. The gas turbine engine as recited in claim 2 , wherein the fan diameter Dfan is greater than claim 2 , or equal to claim 2 , about 73 inches.4. The gas turbine engine as recited in claim 1 , wherein the high pressure turbine section includes two turbine stages.5. The gas turbine engine as recited in claim 1 , wherein a first ratio of a number of stages of the low pressure compressor section to a number of stages of the high pressure compressor section is greater than claim 1 , or equal to claim 1 , about 1.6.6. The gas turbine engine as recited in claim 5 , wherein the low pressure turbine section ...

Rotor bow management

A method of reducing rotor bow in a high pressure rotor of a gas turbine engine that has in axial flow a low pressure rotor and a high pressure rotor. The method involves storing bleed air from the gas turbine engine when the engine is running to provide stored pneumatic energy; and using that stored pneumatic energy after the engine has been shut-down to rotate the high pressure rotor at a speed and for a duration that reduces rotor bow. A gas turbine engine wherein rotor bow in the high pressure rotor after engine shut-down has been reduced by carrying out the aforesaid method is also disclosed.

BYPASS VALVE ASSEMBLY FOR TURBINE GENERATORS

A bypass valve assembly for a turbine generator includes a valve body, bypass seats, valve stem, valve cap, bypass valve disc, bypass valves, and pressure seal head. The valve body defines a central bore and a plurality of passageways. Each passageway has an inlet smaller than its outlet. Each bypass seat is within the inlet of a corresponding passageway. The bypass seats have a higher wear resistance than the valve body. The valve stem is within the central bore. The valve cap is secured to the valve body. The bypass valve disc is secured to the valve stem. Each bypass valve has a base portion and a nose portion. Each nose portion defines a contoured surface area with a wear coating and extends into a corresponding passageway. The pressure seal head is disposed around the valve stem and defines steps having a wear coating. 1. A bypass valve assembly for use in a turbine generator , the bypass valve assembly comprising:a valve body defining a central bore and a plurality of passageways, each passageway having a smaller area at an inlet portion and a larger area at an outlet portion to define a flared passageway;a plurality of bypass seats, each bypass seat disposed within the inlet portion of a corresponding passageway of the plurality of passageways, the bypass seats being formed of a material having higher wear resistance than the valve body;a valve stem disposed within the central bore of the valve body;a valve cap secured to a distal portion of the valve body;a bypass valve disc secured to a distal end portion of the valve stem;a plurality of bypass valves, each bypass valve having a base portion and a nose portion, each nose portion defining a predefined contoured surface area and extending into a corresponding passageway of the plurality of passageways, and at least a portion of the contoured surface area having a wear coating disposed thereon; anda pressure seal head disposed around the valve stem, the pressure seal head defining steps having a wear coating ...

Jet engine cold air cooling system

Methods and devices for cooling systems ( 100, 700 ) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling systems include: a first precooler ( 210 ) receiving bleed air from the jet engine compressor; a heat exchanger ( 730 ) downstream from the first precooler ( 210 ); a cooling system compressor ( 220 ) downstream from the first precooler ( 210 ), wherein the heat exchanger ( 730 ) and the cooling system compressor ( 220 ) are in separate flow paths from the first precooler ( 210 ); a cooling system precooler ( 230 ) downstream from the cooling system compressor ( 220 ); a cooling system turbine ( 240 ) with variable guide vanes—VGT—and downstream from the cooling system precooler ( 230 ); and a discharge conduit ( 245 ) downstream from the cooling system turbine ( 240 ) and the heat exchanger ( 730 ). A bypass line ( 290 ) can also be included that bypasses the cooling system turbine ( 240 ).

GAS TURBINE WITH AN AIR BLEEDER TUBE

A gas turbine provided with an air bleeder tube () that, during startup, bleeds a portion of the compressed air of a compressor from the compressor and discharged the bled air into a cylindrical exhaust duct (), wherein the air bleeder tube () is disposed at a portion that does not obstruct the flow of the main flow of combustion gas. 1said exhaust duct is a cylindrical shape and a rectangular exhaust duct of which a rectangular cross section is connected to a lower end of said exhaust duct and said air bleeder tube is arranged at a corner of an inlet end of said rectangular exhaust duct.. A gas turbine comprising a compressor and an air bleeder tube, wherein a portion of compressed air is bled from said compressor at a startup timing and said compressed air is discharged into an exhaust duct through said air bleeder tube, said gas turbine characterized in that a main flow of combustion gas is not obstructed by an arrangement of said air bleeder tube, and This application is a divisional of copending U.S. patent application Ser. No. 13/977,172, filed on Oct. 28, 2013, and wherein U.S. patent application Ser. No. 13/977,172 is a National Stage application filed under 35 U.S.C. §371 of International Application No. PCT/JP2012/052084, filed on Jan. 31, 2012, and which is based upon and claims the benefit of priority under 35 U.S.C. §119(a) of Japanese Patent Application No. 2011-039204, filed on Feb. 25, 2011, the entire contents of which are incorporated herein by reference.The present invention relates to a gas turbine.In the conventional gas turbine, it has been known a portion of compressed air bled from a compressor and a bled air is discharged into an exhaust duct through an air bleeder tube (for example, Patent Document 1 as described below).is a schematic view for showing one example of a conventional gas turbine. As shown in , the conventional gas turbine comprises a rotor for rotating around a rotational axis, a plurality of rotor blades circularly mounted at ...

Method for the Calculation of the Working Fluid Loss in an Organic Rankine Cycle Plant

Method for the calculation of the working fluid loss in an organic Rankine cycle plant, comprising at least one evaporator (), a preheater (), a turbine (), a condenser (), a pump (), a collecting well () and a process piping (), wherein said working fluid, when the plant is stopped, is in part present in known volumes inside the plant and partly drained in at least a storage tank () comprising at least three rooms or volumes: —a first volume (Vck) for storing the fluid to be measured, —a second volume (Vc) having a restricted section for measuring the volume of fluid stored in said first volume (Vck), —a third volume (Vckd) containing the portion of the fluid already measured, wherein, in said method, the working fluid loss of the plant is calculated as the difference between the fluid amount measured in two different instants of time. 2. The method according to claim 1 , wherein said plant is realized by the following phases:a) activation of a bypass of the hot spring{'b': 12', '2, 'b) activation of a bypass () of the turbine ()'}{'b': 5', '1', '4', '4', '13', '6, 'c) filling of the preheater () and the evaporator () up to a known level, stopping of the pump () and closing a shut-off valve downstream of the pump () d) opening of the exhaust manifolds of the drain valves () of the air condensers (ACC) for the drainage of the liquid in the volume (Vck) of the tank storage ().'}34. The method according to wherein said step b) further comprises the ramp reduction of the pump load () until the normal plant stop in a time not exceeding 30 min.412211. The method according to claim 2 , wherein said step c) is characterized by the simultaneous closing of the by-pass () of the turbine () and the inlet valves ().5. The method according to claim 2 , wherein said step c) further comprises the achievement of balance between the working fluid and the ambient temperature in a time not exceeding 30 min.644. The method according to claim 2 , wherein said step c) of pump () stopping ...

Selectable Barrier Filtration System

An air intake system (AIS) has a plenum and an inlet barrier filter associated with the plenum, through which air can selectively enter the plenum. The AIS also has an inlet duct associated with the plenum, through which air can selectively enter the plenum. The AIS also has a bypass door associated with the inlet duct, the bypass door being configured to selectively change an amount of air allowed to pass through the inlet duct. The AIS also has a filter airflow change device configured to change an amount of airflow allowed through the inlet barrier filter.

DUAL-FED AIRFOIL TIP

An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity. 1. An airfoil of a gas turbine engine comprising:a leading edge extending in a radial direction;a tip extending in an axial direction from the leading edge;a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil;a second rib, the second rib and the first rib defining a serpentine cavity therein;a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity; anda bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.2. The airfoil of claim 1 , further comprising a divider portion located proximate to the bypass aperture within the serpentine cavity and configured to aid in directing (i) a first portion of air from the serpentine cavity into the flag tip cavity and (ii) a second portion of air within the serpentine cavity.3. The airfoil of claim 2 , wherein the divider portion is connected to the third rib.4. The airfoil of claim 2 , wherein the divider portion extends a predetermined length into the serpentine cavity.5. The airfoil of ...

Aircraft comprising a turbine engine incorporated into the rear fuselage with variable supply

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans ( 7, 8 ), the turbine engine being incorporated at the rear of a fuselage ( 1 ) of the aircraft, in the extension of same and comprising at least two gas generators ( 2 a, 2 b ) that supply, via a shared central stream ( 4 ), a power turbine ( 3 ), the turbine ( 3 ) comprising two contrarotating rotors ( 5, 6 ) for driving two fans ( 7,8 ) disposed downstream from the gas generators ( 2 a, 2 b ), said aircraft comprising means ( 15 ) arranged for separating the gas flow in the power turbine ( 3 ) into at least two concentric streams ( 16, 17 ) and a device comprising first means for distributing the gas flow ( 21 - 24 ) between said streams ( 16, 17 ) from the central stream ( 4 ), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream ( 16 ) of the streams ( 16, 17 ) of the power turbine ( 3 ).

METHOD FOR CONTROLLING SURGE MARGIN OF GAS TURBINE AND EXTRACTION DEVICE FOR GAS TURBINE

Disclosed herein is a method for controlling a surge margin of a gas turbine and an extraction device for a gas turbine. The method for controlling the surge margin of the gas turbine and the extraction device for the gas turbine may support stable operation of a compressor unit in the gas turbine, thereby improving the efficiency of the gas turbine and minimizing vibration and noise of the gas turbine. 1. A method for controlling a surge margin of a gas turbine , comprising:determining a number of revolutions and pressure ratios of compressors located at an initial stage among a plurality of compressor stages installed in a compressor unit of a gas turbine; andcontrolling a surge margin by extracting compressed air from each of the compressor stages to the gas turbine, when the surge margin at the initial compressor stage does not satisfy a reference margin.2. The method of claim 1 , wherein the controlling of the surge margin comprises a first extraction step of extracting the compressed air from a first compressor in which compression is initially performed.3. The method of claim 2 , wherein in the first extraction step claim 2 , an amount of extracted air supplied to the gas turbine changes depending on a state of the surge margin.4. The method of claim 2 , wherein the controlling of the surge margin further comprises a second extraction step of additionally extracting the compressed air from a neighboring compressor stage located at a next stage of the first compressor claim 2 , when the surge margin is not stabilized after the first extraction step.5. The method of claim 4 , wherein in the second extraction step claim 4 , the amount of extracted air supplied to the gas turbine changes depending on the state of the surge margin.6. The method of claim 4 , wherein the controlling of the surge margin further comprises a third extraction step of additionally extracting the compressed air from a next compressor stage claim 4 , when the surge margin is not stabilized ...

GAS TURBINE ENGINE FLUID COOLING SYSTEMS AND METHODS OF ASSEMBLING THE SAME

A fluid cooling system for use in a gas turbine engine including a fan casing circumscribing a core gas turbine engine includes a heat source configured to transfer heat to a heat transfer fluid and a primary heat exchanger coupled in flow communication with the heat source. The primary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the fan casing. The fluid cooling system also includes a secondary heat exchanger coupled in flow communication with the primary heat exchanger. The secondary heat exchanger is configured to channel the heat transfer fluid therethrough and is coupled to the core gas turbine engine. The fluid cooling system also includes a bypass mechanism coupled in flow communication with the secondary heat exchanger. The bypass mechanism is selectively moveable based on a temperature of a fluid medium to control a cooling airflow through the secondary heat exchanger. 1. A fluid cooling system for use in a gas turbine engine including a core section of the gas turbine engine having an axis of rotation and a fan casing substantially circumscribing the core section of the gas turbine engine , said fluid cooling system comprising:a heat source configured to transfer heat to a heat transfer fluid;a primary heat exchanger coupled in flow communication with said heat source and configured to channel the heat transfer fluid therethrough, said primary heat exchanger coupled to the fan casing;a secondary heat exchanger coupled in flow communication with said primary heat exchanger and configured to channel the heat transfer fluid therethrough, said secondary heat exchanger located within to the core section of the gas turbine engine;a bypass mechanism coupled in flow communication with said secondary heat exchanger, said bypass mechanism being selectively moveable based on a temperature of a fluid medium to control a cooling airflow through said secondary heat exchanger.2. The fluid cooling system in accordance ...

HEAT EXCHANGER SYSTEM FOR AIRCRAFT AND ASSOCIATED METHOD OF OPERATION

The heat exchanger system can have a first conduit extending from at least one first conduit inlet through a heat exchanger to at least two first conduit outlets; a second conduit extending from at least one second inlet through the heat exchanger to at least one second outlet, the first and second conduits disposed adjacent to one another in heat exchange engagement within the heat exchanger; and a bypass conduit extending from the first conduit between the at least one first inlet and the heat exchanger to the first conduit between the heat exchanger and at least one of said at least two first outlets. 1. A gas turbine engine comprising:a first conduit extending from at least one first conduit inlet through a heat exchanger to at least two first conduit outlets;a second conduit extending from at least one second inlet through the heat exchanger to at least one second outlet, the first and second conduits disposed adjacent to one another in heat exchange engagement within the heat exchanger; anda bypass conduit extending from the first conduit between the at least one first inlet and the heat exchanger to the first conduit between the heat exchanger and at least one of said at least two first outlets.2. The gas turbine engine of wherein the bypass conduit has a passive flow meter.3. The gas turbine engine of wherein the bypass conduit has a control valve.4. The gas turbine engine of wherein the at least two first conduit outlets include a first outlet and a second outlet claim 1 , and the first conduit has a first outlet branch extending between the heat exchanger and the first outlet claim 1 , and a second outlet branch extending between the heat exchanger and the second outlet claim 1 , wherein the bypass conduit is connected to the second outlet branch.5. The gas turbine engine of further comprising an auxiliary branch extending between the second outlet branch and the first outlet branch.6. The gas turbine engine of further comprising a control valve in at ...