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Небесная энциклопедия

Космические корабли и станции, автоматические КА и методы их проектирования, бортовые комплексы управления, системы и средства жизнеобеспечения, особенности технологии производства ракетно-космических систем

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Мониторинг СМИ

Мониторинг СМИ и социальных сетей. Сканирование интернета, новостных сайтов, специализированных контентных площадок на базе мессенджеров. Гибкие настройки фильтров и первоначальных источников.

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 4312. Отображено 100.
09-02-2012 дата публикации

Ventilation inlet

Номер: US20120034068A1
Автор: Zahid M. Hussain
Принадлежит: Rolls Royce PLC

A ventilation inlet comprising a ventilation pipe to receive flow from a first flow zone and to deliver the flow to a second flow zone; a divider arranged to divide a portion of the ventilation pipe into a static pressure zone and a total pressure zone; and a deflector arranged to direct flow from the total pressure zone at least partially across the static pressure zone to restrict delivery of the flow from the static pressure zone to the second flow zone dependent on the pressure of the flow in the first flow zone.

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Номер записи: 1
09-02-2012 дата публикации

Actuation mechanism for a convertible gas turbine propulsion system

Номер: US20120034080A1
Автор: Gregory E. Reinhardt, Rajendra K. Agrawal
Принадлежит: Agrawal Rajendra K, Reinhardt Gregory E

A method of powering a rotary-wing aircraft includes selectively coupling and uncoupling a first power turbine to change a power distribution between the rotor system and the secondary propulsion system.

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Номер записи: 2
10-05-2012 дата публикации

Variable area fan nozzle fan flutter management system

Номер: US20120110980A1
Автор: Robert J. Morris, Zbigniew M. Grabowski
Принадлежит: Individual

A system and method of controlling a fan blade flutter characteristic of a gas turbine engine includes adjusting a variable area fan nozzle in response to a neural network.

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Номер записи: 3
10-05-2012 дата публикации

Control method for cooling a turbine stage in a gas turbine

Номер: US20120111020A1
Автор: Federico Bonzani, Luca Bozzi, Marco Mantero
Принадлежит: Ansaldo Energia SpA

A control method for cooling a turbine stage of a gas turbine, whereby cooling air is bled from combustion air flowing in a compressor of the gas turbine, and is fed to a cooling circuit staring from a stator of the turbine stage; and cooling airflow is adjusted as a function of the pressure at the inlet of the cooling circuit, and as a function of the combustion air pressure at the exhaust of the compressor; more specifically, there is a feedback control setting a setpoint, which is predetermined as a function of the power output of the turbine to reduce contaminating emissions.

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Номер записи: 4
21-06-2012 дата публикации

Hot gas path component cooling for hybrid pulse detonation combustion systems

Номер: US20120151896A1
Автор: Adam Rasheed, Brian Gene Brzek, Douglas Carl Hofer, Fuhua Ma, Narendra Digamber Joshi, Thomas Michael Lavertu, Venkat Eswarlu Tangirala
Принадлежит: General Electric Co

The flow through the core of a hybrid pulse detonation combustion system is passed through a compressor and then separated into a primary flow, that passes directly to the combustor, and a bypass flow, which is routed to a portion of the system to be used to cool components of the system. The bypass flow is routed to a nozzle of the pulse detonation combustor. The flow is then passed back into the primary flow through the core downstream of where it was extracted.

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Номер записи: 5
21-06-2012 дата публикации

Turbojet including an automatically variable flow rate bleed circuit for cooling air

Номер: US20120151936A1
Автор: Amaury Roussely-Rousseau, Emmanuel Berche, Guy Burlet
Принадлежит: SNECMA SAS

Bleeding cooling air to cool a subassembly, e.g. such as a turbine, with automatic adjustment of the air flow section as a function of the speed of the engine. According to the invention, a shutter element is fastened to co-operate with a bleed hole, with the material that constitutes either the shutter element or the wall in which the hole is formed being of a type in which it is possible to create eddy currents, and a magnet is mounted to move past said arrangement.

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Номер записи: 6
04-10-2012 дата публикации

System and method for air extraction from gas turbine engines

Номер: US20120247113A1
Автор: Michael Joseph Alexander, Paul William Plummer, Timothy Lee Janssen, Vijay Anand Raghavendran Nenmeni
Принадлежит: General Electric Co

The disclosed embodiments relate to a system and method that allows air to be extracted from a plurality of gas turbine engines and fed to a downstream process, even in situations in which one or more of the gas turbine engines are operating in a part load condition. For example, in an embodiment, a method includes monitoring signals representative of a header pressure of a header, or a pressure of extraction air flow from one or more gas turbine engines to the header, or both, and maintaining substantially continuous flows of extraction air from the gas turbine engines to the header. The substantially continuous flows are maintained when the gas turbine engines are under symmetric and asymmetric load conditions.

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Номер записи: 7
01-11-2012 дата публикации

Multiple core variable cycle gas turbine engine and method of operation

Номер: US20120272656A1
Автор: James W. Norris
Принадлежит: United Technologies Corp

A gas turbine engine system includes a fan assembly, a low pressure compressor, a low pressure turbine, a plurality of engine cores including a first engine core and a second engine core, and a control assembly. A primary flowpath is defined through the fan assembly, the low pressure compressor, the low pressure turbine, and the active engine cores. Each engine core includes a high pressure compressor, a combustor downstream from the high pressure compressor, and a high pressure turbine downstream from the combustor. The control assembly is configured to control operation of the plurality of engine cores such that in a first operational mode the first and the second engine cores are active to generate combustion products and in a second operational mode the first engine core is active to generate combustion products while the second engine core is idle.

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Номер записи: 8
17-01-2013 дата публикации

Compressors with integrated secondary air flow systems

Номер: US20130017066A1
Автор: Jong Lee, Nick Nolcheff
Принадлежит: Honeywell International Inc

A compressor includes a rotor platform; a rotor blade; and a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap. A secondary air flow system includes a bleed inlet configured to remove secondary air flow from the primary air flow; an injection opening disposed in the inner surface of the casing upstream of the bleed inlet; an accessory conduit; and a plenum fluidly coupled to the bleed inlet, the injection opening, and the accessory conduit. The bleed inlet and plenum at least partially define a secondary air flow path such that a first portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the injection opening and a second portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the accessory conduit.

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Номер записи: 9
21-02-2013 дата публикации

Control of load rejection

Номер: US20130043680A1
Автор: Martin Liebau, Peter Marx, Sami Hashmi
Принадлежит: Alstom Technology AG

A control system is provided for a power generating system having a gas turbine, a flue gas exhaust stage and a blow-off valve assembly. The gas turbine includes a compression stage, a combustion stage and a driveshaft. The blow-off valve assembly is configured to selectively provide fluid communication between the combustion stage and the flue gas exhaust stage. The control system includes a controller configured to output a signal causing the blow-off valve assembly to provide the fluid communication in response to a sudden de-loading of the gas turbine.

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Номер записи: 10
02-05-2013 дата публикации

Gas turbine engine with two-spool fan and variable vane turbine

Номер: US20130104521A1
Автор: Daniel B. Kupratis
Принадлежит: Individual

A gas turbine engine and a method of operating the gas turbine engine according to an exemplary aspect of the present disclosure includes modulating a variable high pressure turbine inlet guide vane of a high pressure spool to performance match a first stage fan section of a low pressure spool and an intermediate stage fan section of an intermediate spool to maintain a generally constant engine inlet flow while varying engine thrust.

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Номер записи: 11
23-05-2013 дата публикации

EMISSIONS CONTROL SYSTEMS AND METHODS

Номер: US20130125524A1
Автор: Detar Gail Leroy, Heverley, III Lawrence, Kiran Shashi, Kulkarni Shridhar Shrikant, Plummer Jessica Lynn
Принадлежит:

Methods and systems are provided related to an emissions control system. The emissions control system has an exhaust after-treatment system defining a plurality of distinct exhaust flow passages through which at least a portion of an exhaust stream can flow, e.g., the exhaust stream is produced by an engine. The emissions control system also includes a controller for controlling injection of reductant into the exhaust stream flowing through each of the flow passages. In one example, the emissions control system is configured for use in a vehicle, such as a locomotive or other rail vehicle. 1. An emissions control system , comprising:an exhaust after-treatment system defining a plurality of distinct exhaust flow passages through which an exhaust stream can flow, the exhaust stream produced by an engine,wherein the exhaust after-treatment system includes a plurality of first exhaust after-treatment components, each in or otherwise associated with a respective one of the exhaust flow passages.2. The emissions control system of claim 1 , wherein the plurality of exhaust flow passages are positioned at least generally parallel to each other.3. The emissions control system of claim 1 , wherein each of the plurality of exhaust flow passages is defined by a respective substrate claim 1 , or a set of substrates claim 1 , though which the exhaust stream can flow.4. The emissions control system of claim 3 , wherein the substrate claim 3 , or each of the set of substrates claim 3 , is rectangular shaped or cylindrically shaped.5. The emissions control system of claim 3 , wherein at least one of the exhaust flow passages is defined at least in part by a respective set of parallel-arrayed distinct substrates that are arranged with at least one of the parallel-arrayed distinct substrates located on an upper level above at least one other of the parallel-arrayed distinct substrates located on a lower level.6. The emissions control system of claim 5 , wherein a first of the exhaust ...

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Номер записи: 12
23-05-2013 дата публикации

Method and apparatus for optimizing the operation of a turbine system under flexible loads

Номер: US20130125557A1
Автор: Alston I. Scipio, Douglas Corbin Warwick, Joseph Klosinski, Michael Anthony Cocca, Mike Yarnold, Ryan Eric Obenhoff, Sanji Ekanayake, Thomas John Freeman
Принадлежит: Individual

A gas turbine system includes a compressor protection subsystem; a hibernation mode subsystem; and a control subsystem that controls the compressor subsystem and the hibernation subsystem. At partial loads on the turbine system, the compressor protection subsystem maintains an air flow through a compressor at an airflow coefficient for the partial load above a minimum flow rate coefficient where aeromechanical stresses occur in the compressor. The air fuel ratio in a combustor is maintained where exhaust gas emission components from the turbine are maintained below a predetermined component emission level while operating at partial loads.

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Номер записи: 13
06-06-2013 дата публикации

ON BOARD INERT GAS GENERATION SYSTEM

Номер: US20130139521A1
Автор: Das Alok, Joshi Mahesh Prabhakar, Massey Alan Ernest
Принадлежит: EATON AEROSPACE LIMITED

An on board inert gas generation system for an aircraft receives air from a relatively low pressure source such as low pressure engine bleed air or ram air and passes it to a positive displacement rotary compressor to increase the pressure thereof to be suitable for supply to an air separation module. The speed of the positive displacement compressor may be adjusted across a wide range in order to provide efficient operation in cruise and descent phases of aircraft flight. 1. An on board inert gas generation system for use in an aircraft having an on board source of low pressure air , the gas generation system comprising:a positive displacement compressor having an inlet for receiving a portion of the low pressure fluid, and an outlet in flow communication with an air separation module.2. The on board inert gas generation system according to claim 1 , wherein the low pressure air is at least one of:low pressure engine bleed air supplied from a gas turbine power plant on board the aircraft, orram air obtained from a ram air inlet at or adjacent an exterior of the aircraft.3. The on board inert gas generation system according to claim 1 , further comprising an electric motor drivably connected to the positive displacement compressor.4. The on board inert gas generation system according to claim 3 , wherein the electric motor is connectable to receive electrical energy from at least one of: a generator or an energy storage arrangement associated therewith claim 3 , and an aircraft electrical supply.5. The on board inert gas generation system according to claim 4 , further comprising a power controller configured to at least one of:receive electrical energy from the generator or an electrical storage arrangement associated therewith; receive electrical energy from the aircraft electrical supply; or controllably supply electrical energy to the electric motor.6. The on board inert gas generation system according to claim 1 , further comprising a heat exchanger in a flow ...

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Номер записи: 14
06-06-2013 дата публикации

METHOD FOR OPERATING AN EXHAUST GAS SYSTEM OF AN INTERNAL COMBUSTION ENGINE

Номер: US20130144505A1
Автор: Handler Torsten, Neumann Joerg, Pfister Tobias, Wendt Hauke
Принадлежит: ROBERT BOSCH GMBH

A method for operating an exhaust gas system of an internal combustion engine is described, wherein NOx is reduced by means of a SCR catalyst and a NOx reduction capability of an aqueous urea solution to be introduced into the exhaust gas system is monitored, and wherein at least one first variable characterizing the ammonia content of the water is ascertained and an ageing of the aqueous urea solution is inferred from said first variable. 1. A method for operating an exhaust gas system of an internal combustion engine , wherein NOx is reduced by means of a SCR catalyst and a NOx reduction capability of an aqueous urea solution to be introduced into the exhaust gas system is monitored , the method comprising:ascertaining at least one first variable characterizing the ammonia content of the water; andmonitoring the NOx reduction capability of the aqueous urea solution.2. The method according to claim 1 , wherein at least one second variable characterizing the composition of the aqueous urea solution is ascertained and the NOx reduction capability of said aqueous urea solution is inferred from the first and the second variable.3. The method according to claim 1 , wherein the first variable is an electrical conductivity of the aqueous urea solution.4. The method according to claim 1 , wherein the second variable is a density and/or a refractive index and/or a sound velocity and/or a thermal conductivity and/or a dielectric permittivity of the aqueous urea solution.5. The method according to claim 1 , wherein the ascertained first and second variables are linked to one another by means of at least one characteristic diagram and/or a table and/or a mathematical formula in order to infer the NOx reduction capability of the aqueous urea solution.6. The method according to claim 1 , wherein a filling level of a reservoir of the reducing agent and/or a point in time of a filling of the reservoir and/or a temperature of the reducing agent are used in a complementary manner ...

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Номер записи: 15
13-06-2013 дата публикации

Gas turbine engine with fan variable area nozzle for low fan pressure ratio

Номер: US20130145745A1
Автор: Frederick M. Schwarz, Gregory A. Kohlenberg, Sean P. Zamora
Принадлежит: Individual

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45.

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Номер записи: 16
27-06-2013 дата публикации

Method for Operating a Combined Cycle Power Plant

Номер: US20130160424A1
Автор: Broesamle Stefan Eduard, Ruchti Christoph
Принадлежит: ALSTOM TECHNOLOGY LTD.

In a method for operating a combined cycle power plant (), which has a gas turbine installation () and a water-steam cycle () connected to the gas turbine installation () by a waste heat steam generator () and has at least one steam turbine (), the gas turbine installation () includes a compressor (), a combustion chamber (), and a turbine (). To cool the turbine (), air compressed at the compressor () is removed, cooled in at least one cooler () flowed through by water, thus generating steam, and introduced into the turbine (). At least with the gas turbine installation () running, prior to or during the start-up of the water-steam cycle (), waste heat, which is contained in the steam generated in the at least one cooler (), is used to good effect for pre-heating the installation inside the combined cycle power plant (). 1. A method for operating a combined cycle power plant , which plant has a gas turbine installation and a water-steam cycle , which water-steam cycle is connected to the gas turbine installation by a waste heat steam generator and has at least one steam turbine , wherein the gas turbine installation includes a compressor , a combustion chamber , and a turbine , and wherein , to cool the turbine , air compressed at the compressor is removed , cooled in at least one cooler flowed through by water thus generating steam , and introduced into the turbine , the method comprising:at least with the gas turbine installation running, and prior to or during the start-up of the water-steam cycle, pre-heating at least a portion of the installation at a position inside the combined cycle power plant with waste heat contained in the steam generated in the at least one cooler.2. A method according to claim 1 , wherein the combined cycle power plant includes means for blowing-off or condensing steam generated in the at least one cooler.3. A method according to claim 2 , further comprising:blowing-off steam generated in the at least one cooler via a secondary stack. ...

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Номер записи: 17
04-07-2013 дата публикации

Combustor system for use in turbine engines and methods of operating a turbine engine

Номер: US20130167545A1
Автор: Ahmed Mostafa Elkady, Christian Lee Vandervort, Narendra Joshi, Sherif Mohamed
Принадлежит: General Electric Co

A combustor system for use in a turbine engine is provided. The turbine engine includes turbine assembly that includes a fluid inlet, a fluid outlet, and a combustion gas path defined therebetween. The combustor system includes a first combustor assembly and a second combustor assembly. The first combustor assembly is coupled to the turbine assembly for channeling a first flow of combustion gases through the turbine assembly. The first combustor assembly is oriented adjacent to the turbine assembly inlet to channel the first flow of combustion gases to the turbine assembly through the turbine assembly inlet. The second combustor assembly is coupled to the turbine assembly along the combustion gas path for channeling a second flow of combustion gases through the turbine assembly.

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Номер записи: 18
11-07-2013 дата публикации

Flushing the exhaust gas recirculation lines of a gas turbine

Номер: US20130174535A1
Автор: Floris Van Straaten, Juergen Hoffmann
Принадлежит: Alstom Technology AG

A method and gas turbine are provided for the reliable purging of an exhaust gas recirculation line of the gas turbine with exhaust gas recirculation without the use of additional blow-off fans. A blow-off flow of the compressor is used for the purging of the exhaust gas recirculation line. The gas turbine can include at least one purging line which connects a compressor blow-off point to the exhaust gas recirculation line.

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Номер записи: 19
25-07-2013 дата публикации

Method and device for detecting a rotational separation adversely affecting a turbine engine compressor

Номер: US20130186191A1
Автор: Cedrik Djelassi
Принадлежит: SNECMA SAS

A method and device detecting a rotating stall affecting a turbine engine compressor. The detection method includes: detecting an abnormal acceleration of the engine or an operating line of the compressor that is characteristic of a failure of the engine; storing a reference temperature measured at an outlet from a turbine of the engine at an instant of detection; comparing a determined temperature threshold with the difference between a current temperature at the outlet of the turbine as measured after detection and the reference temperature; and identifying that a rotating stall is present in event of the threshold being exceeded.

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Номер записи: 20
01-08-2013 дата публикации

Buffer system that communicates buffer supply air to one or more portions of a gas turbine engine

Номер: US20130192251A1
Автор: Peter M. Munsell, Philip S. Stripinis
Принадлежит: United Technologies Corp

A gas turbine engine includes a buffer system that can communicate a buffer supply air to a portion of the gas turbine engine. The buffer system includes a first bleed air supply having a first pressure, a second bleed air supply having a second pressure that is greater than the first pressure, and an ejector that selectively augments the first bleed air supply to prepare the buffer supply air for communication to the portion of the gas turbine engine.

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Номер записи: 21
31-10-2013 дата публикации

ROTARY VANE ACTUATOR OPERATED AIR VALVES

Номер: US20130283762A1
Автор: Mailander William James, Moster David Anthony, Plaatje Matthew John, Simpson Andrew David
Принадлежит: GENERAL ELECTRIC COMPANY

Rotary vane actuator operated air valves associated with gas turbine engines are disclosed. An example gas turbine engine may include a fan, a compressor, a combustor, and a turbine in a serial flow relationship; a supply pipe arranged to convey compressed air from one or more of the fan and the compressor; a valve operatively disposed in the supply pipe, the valve including a rotatable valve member arranged to modulate flow of the compressed air through the supply pipe based upon an angular position of the valve member, the valve member being rotatable between an open position and a shut position; and/or a hydraulically operated rotary vane actuator operatively coupled to rotate the valve member. 1. A gas turbine engine comprising:a fan, a compressor, a combustor, and a turbine in a serial flow relationship;a supply pipe arranged to convey compressed air from one or more of the fan and the compressor;a valve operatively disposed in the supply pipe, the valve comprising a rotatable valve member arranged to modulate flow of the compressed air through the supply pipe based upon an angular position of the valve member, the valve member being rotatable between an open position and a shut position; anda hydraulically operated rotary vane actuator operatively coupled to rotate the valve member.2. The gas turbine engine of claim 1 , wherein the gas turbine engine is arranged to provide propulsion for an aircraft in flight.3. The gas turbine engine of claim 1 , wherein the rotary vane actuator is hydraulically operated by pressurized fuel.4. The gas turbine engine of claim 1 ,wherein the turbine comprises a high pressure turbine; andwherein the supply pipe is arranged to convey the compressed air from the fan to a high pressure turbine active clearance control system.5. The gas turbine engine of claim 1 ,wherein the turbine comprises a low pressure turbine; andwherein the supply pipe is arranged to convey the compressed air from the fan to a low pressure turbine active ...

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Номер записи: 22
07-11-2013 дата публикации

COMBUSTOR MIXING JOINT AND METHODS OF IMPROVING DURABILITY OF A FIRST STAGE BUCKET OF A TURBINE

Номер: US20130291548A1
Автор: Ingram Clint L., Siden Gunnar Leif
Принадлежит:

The present application and the resultant patent provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface; passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region; substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine. 1. A method of improving durability of a first stage bucket of a turbine of a gas turbine engine , the method comprising:generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface;passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region;substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; andpassing the mixed combustion flow to a first stage bucket of a turbine.2. The method of claim 1 , wherein passing the mixed combustion flow to the first stage bucket comprises generating a substantially uniform velocity field in the first stage bucket.3. The method of claim 1 , wherein passing the mixed combustion flow to the first stage bucket comprises generating a substantially uniform temperature field in the first stage bucket.4. The method of claim 1 , wherein the first combustion flow and the second combustion flow are passed to the mixing region at a first velocity claim 1 , wherein the mixed combustion flow is passed to the ...

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Номер записи: 23
26-12-2013 дата публикации

Spherical-link end damper system with near constant engagement

Номер: US20130343876A1
Автор: Kevin J. Cummings, Ryan Edward LeBlanc
Принадлежит: United Technologies Corp

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.

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Номер записи: 24
02-01-2014 дата публикации

Scheduling of variable area fan nozzle to optimize engine performance

Номер: US20140005910A1
Автор: Christopher J. Hanlon, Glenn D. Bartkowski, John R. Gendron, Matthew D. Teicholz, Paul H. Spiesman
Принадлежит: Individual

A disclosed control system for a gas turbine engine includes a controller configured to set a position of the variable area fan nozzle according to a predetermined schedule of variable area fan nozzle positions corresponding to a flight operating condition. The schedule is determined in view of a relationship between a position of the variable area nozzle and a performance level of the engine at current flight conditions.

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Номер записи: 25
09-01-2014 дата публикации

Aircraft Maintenance Apparatus

Номер: US20140007587A1
Автор: Borgess Jason, Grebe Wayne, Heydinger Peter, Spinazze Paul A., Wallace Rodney
Принадлежит: Tronair Inc.

This Mobile Driven Hydraulic Power Unit (HPU) provides a source of clean, pressurized hydraulic fluid for performing required aircraft maintenance. The unit sits on a heavy duty towable frame designed for off-road terrain. Electric start engine, fully instrumented operation panel for both the engine and hydraulic system. A gasoline engine or diesel engine may be used to power the HPU. 1. An apparatus for servicing an aircraft , comprising a mobile vehicle;a hydraulic power unit (HPU) detachably mounted to the vehicle;the HPU further comprising a computer; andthe HPU being capable of providing hydraulic fluid and outputs.2. An apparatus according to claim 1 , wherein the HPU further comprises:a hydraulic pump for producing a fluid hydraulic pressure output to the aircraft.3. An apparatus according to further comprising outputs for pressurized hydraulic fluid.4. An apparatus according to further comprising a gasoline engine for driving the HPU.5. An apparatus according to further comprising a diesel engine for driving the HPU.6. An apparatus according to capable of providing hydraulic fluid at a hydraulic pressure ranging from 250-1500 psi.7. An apparatus according to capable of providing hydraulic fluid at a hydraulic pressure ranging from 250 to 4000 psi. The present patent application is based upon and claims the benefit of provisional patent application No. 61/668,166, filed on Jul. 5, 2012.The present invention relates to the field of aviation and, more particularly, to the maintenance of aircraft.Various aircraft maintenance equipment has been developed for maintaining various portions of an aircraft. Aircraft ground servicing, specifically, provides electrical, hydraulic fluid, and gaseous inputs to aircraft at or on remote locations. An aircraft on the ground whose engine is not functioning requires a number of services to determine whether the aircraft is in a condition to fly or taxi. These services include: electrical power, hydraulic power, engine-start ...

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Номер записи: 26
06-02-2014 дата публикации

Flow discharge device

Номер: US20140033733A1
Автор: Kevin Mark Britchford, Pamela Joan PHILLIPSON, Vincent Benoit SAVARIN
Принадлежит: Rolls Royce PLC

A bleed flow discharge device ( 136 ) adapted to discharge a bleed fluid flow into a main fluid flow, wherein the bleed flow discharge device comprises an outer wall ( 135 ) defining a passage ( 137 ) for the bleed fluid flow, the outer wall comprising a wave-shaped edge ( 139 ) where the bleed fluid flow meets the main fluid flow.

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Номер записи: 27
27-02-2014 дата публикации

Nacelle scoop inlet

Номер: US20140053532A1
Автор: Steven H. Zysman
Принадлежит: Individual

A scoop inlet for use in a gas turbine engine nacelle has a scoop inlet, and a tab extending forwardly of the scoop inlet. The scoop communicates with a downstream flowpath. The tab has at least one opening at a location upstream of the scoop inlet. A nacelle and a gas turbine engine are also disclosed.

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Номер записи: 28
27-02-2014 дата публикации

GAS TURBINE

Номер: US20140053572A1
Автор: Ishizaka Koichi, Ito Eisaku, Machida Yukinori, Sakamoto Yasuro, Wakazono Susumu
Принадлежит: MITSUBISHI HEAVY INDUSTRIES, LTD.

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. 1. 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.2. A gas turbine as claimed in claim 1 , said gas turbine characterized in that said air bleeder tube is passed through an inside of an structural element connected to a bearing for supporting a rotor.3. A gas turbine as claimed in claim 1 , said gas turbine characterized in that said air bleeder tube is arranged at a portion immediately near a downstream end of a structural element connected to a bearing for supporting a rotor along said main flow of combustion gas.4. A gas turbine as claimed in claim 1 , said gas turbine characterized in that an outlet of said air bleeder tube is arranged at a lower outer peripheral portion of a structural element connected to a bearing for supporting a rotor.5. A gas turbine as claimed in claim 1 , said gas turbine characterized of further comprising three structure elements connected to a bearing for supporting a rotor claim 1 , wherein said three structural elements are arranged in an invert Y shape along a peripheral direction of a rotor and said air bleeder tube is passed through an inside of said structural elements.6. A gas turbine as claimed in claim 1 , said gas turbine characterized in that said exhaust duct is a cylindrical shape and a rectangular exhaust duct of which a rectangular cross section is connected to a lower end ...

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Номер записи: 29
13-03-2014 дата публикации

METHOD FOR OPERATING A THERMAL POWER PLANT

Номер: US20140069104A1
Автор: Chan Kevin, Li Hongtao, Oesterheld Joerg
Принадлежит: ALSTOM Technology Ltd

The invention relates to a method for operating a thermal power plant, which includes a gas turbine and a generator driven directly by the gas turbine by means of a shaft and being connected to an electrical grid having a grid frequency (F) via an electronic decoupling apparatus and a step-up transformer. A synthetic inertia response is achieved by said method includes the steps of: 1. A method for operating a thermal power plant , which includes a gas turbine and a generator driven directly by the gas turbine by means of a shaft and being connected to an electrical grid having a grid frequency (F) via an electronic decoupling apparatus and a step-up transformer , said method comprising the steps of:{'sub': 'G', 'sensing said grid frequency (F);'}{'sub': 'G', 'detecting if in case of an excursion of said grid frequency (F) additional inertial power (ΔP) is required or not;'}if inertial power (ΔP) is required, calculating the magnitude and duration of the additional inertial power (ΔP); andreleasing additional inertial power (ΔP) to said electrical grid in accordance with said calculations via said electronic decoupling apparatus.2. The method according to claim 1 , wherein the detecting step is based on a predefined rate of change of said grid frequency (F) and a predetermined frequency threshold of said grid frequency (F).3. The method according to claim 1 , wherein said electronic decoupling apparatus has a short-term capacity claim 1 , and within said calculating step said short-term capacity of the electronic decoupling apparatus and/or the initial operating conditions of the power plant at start of said grid frequency excursion are considered.4. The method according to claim 1 , wherein for said power releasing step set points are given to said electronic decoupling apparatus to release an active inertial power and also a reactive power to said electrical grid.5. The method according to wherein an upper and lower threshold for said grid frequency (F) or turbine ...

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Номер записи: 30
20-03-2014 дата публикации

EASILY ADAPTABLE COMPRESSOR BLEED SYSTEM DOWNSTREAM OF A VANE PLATFORM

Номер: US20140075955A1
Автор: Twell Philip
Принадлежит: SIEMENS AKTIENGESELLSCHAFT

Discrete bleed behind stator vane platform is provided. The discrete bleed behind stator vane platform relates to a system for bleeding off a working fluid from an inner volume (Vi) of a turbo-machine. The system includes a vane carrier with an annular rail and a vane device comprising at least one vane element, a vane platform and a vane root. The vane element is mounted to the vane platform and the vane root is mounted to the annular rail. A first annular cavity is formed between the vane platform and the annular rail, wherein an annular gap is formed between an edge of the vane platform and the vane carrier such that a part of the working fluid of the turbo-machine is bleedable through the annular gap into the first annular cavity. A second annular cavity is formed between the vane root, the annular rail and the vane carrier, wherein at least one inlet hole is formed into the annular rail for coupling the first annular cavity and the second annular cavity. 1. A system for bleeding off a working fluid from an inner volume (Vi) of a turbo-machine , the system comprisinga vane carrier comprising an annular rail, a vane device comprising at least one vane element, at least one vane platform and a vane root,wherein the vane element is mounted to the vane platform,wherein the vane root is mounted to the annular rail,wherein a first annular cavity is formed between the vane platform and the annular rail,wherein an annular gap is formed between an edge of the vane platform and the vane carrier such that a part of the working fluid of the turbo-machine is bleedable through the annular gap into the first annular cavity,wherein a second annular cavity is formed between the vane root, the annular rail and the vane carrier, andwherein at least one inlet hole is formed into the annular rail for coupling the first annular cavity and the second annular cavity.2. The system according to claim 1 ,wherein the vane element comprises a leading edge and a trailing edge,wherein the ...

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Номер записи: 31
20-03-2014 дата публикации

TURBOMACHINE WITH BLEED VALVES LOCATED AT THE INTERMEDIATE CASE

Номер: US20140075956A1
Автор: PATSOURIS Emmanuel
Принадлежит: SNECMA

An assembly including an intermediate case of a bypass turbojet engine and of an inter-jet case extending upstream of the intermediate case to separate a primary air jet of the turbojet engine from its bypass air jet, the inter-jet case including, passing through it, a closable duct for diverting part of the primary flow to the bypass flow thereby forming a blow-off valve for the LP compressor, the intermediate case including arms passing across the bypass flow and the inter-jet case in its internal cavity including a first chamber situated upstream of the arms and a second chamber situated level with the arms, the duct being open or closed off by an annular component capable of axial movement set in motion by an arm that can rotate about a fixed pivot under action of a control cylinder, the cylinder being positioned in the second chamber. 16-. (canceled)7. An assembly comprising:an intermediate case of a bypass turboduct engine and an inter-duct case extending upstream of the intermediate case to separate a primary air duct of the turboduct engine from a bypass air duct,the inter-duct case including, passing through it, a closable duct for diverting part of a primary flow to the bypass flow, thereby forming a bleed valve for a LP compressor,the intermediate case comprising arms passing through the bypass flow, andthe inter-duct case comprising in its inner cavity a first chamber situated upstream of the arms and a second chamber situated level with the arms,the duct being opened or closed by an annular component capable of axial movement set in motion by an arm that can rotate about a fixed pivot under action of a control cylinder positioned in the second chamber.8. The assembly as claimed in claim 7 , wherein the cylinder is oriented in the circumferential direction of the intermediate case.9. The assembly as claimed in claim 8 , wherein the cylinder is mounted in a manner freely rotatable about an axis oriented radially.10. The assembly as claimed in claim 7 , ...

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Номер записи: 32
06-01-2022 дата публикации

Turbocharger System For A Two-Stroke Engine

Номер: US20220003154A1
Автор: BUCHWITZ James, HANSON Reed, HEDLUND Darren, SALFER Lucas, Trihey Sean, ZIMNEY Derek
Принадлежит:

A turbocharger and method of controlling the same includes a turbine housing comprising an inlet and an outlet, turbine wheel coupled to a shaft. The turbine housing comprising a first scroll and a second scroll for fluidically coupling the inlet and the turbine wheel. The first scroll has a first end adjacent the inlet and a second end adjacent the turbine wheel. The second scroll has a third end adjacent the inlet and a fourth end adjacent the turbine wheel. An exhaust gas diverter valve is coupled to the turbine housing restricting flow into the first scroll or the second scroll. 112-. (canceled)13. A system comprising:a turbocharger comprising a turbine portion and a compressor portion;an engine comprising a throttle body;a boost box;a bypass path coupling the boost box to ambient air outside the boost box;a one way valve coupled in the bypass path communicating air through the one way valve when a first pressure in the boost box is lower than air pressure outside the boost box.14. The system set forth in wherein the bypass path bypasses the compressor portion by providing an alternate air path to the throttle body.15. The system set forth in wherein the bypass path is formed by a duct fluidically coupled to the boost box.16. The system set forth in wherein the duct communicates ambient air to the boost box from an upper plenum.17. The system set forth in wherein the duct communicates ambient air to the boost box from outside the vehicle.18. The system set forth in wherein the duct is formed by an outer wall of a fuel tank.19. The system set forth in wherein the bypass path is disposed on a rearward facing surface of the boost box.2026-. (canceled) This application is a divisional of U.S. patent application Ser. No. 16/691,995, filed Nov. 22, 2019, which claims priority to U.S. Provisional Application No. 62/776,571, filed on Dec. 7, 2018. The above-mentioned patent applications are incorporated herein by reference in its entirety.The present disclosure relates ...

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Номер записи: 33
05-01-2017 дата публикации

AIRCRAFT TURBOMACHINE COMPRISING A HEAT EXCHANGER OF THE PRECOOLER TYPE

Номер: US20170002747A1
Автор: De Vulpillieres Eric, Fert Jérémy Edmond, Pavillet Julien
Принадлежит: SNECMA

The invention relates to an aircraft turbomachine comprising a nacelle and an engine comprising at least one outflowing jet of air, wherein a heat exchanger of the precooler type for supplying air to the aircraft is mounted in the nacelle, said exchanger comprising a primary circuit, the inlet of which is connected to means for taking compressed air from the engine and the outlet of which is connected to means for supplying air to the aircraft, and a secondary circuit supplied with air taken from said air flow. 1. Aircraft turbine engine , comprising a nacelle and an engine having at least one flow duct for an air flow , wherein a heat exchanger of the precooler type for supplying air to the aircraft is mounted in the nacelle , said exchanger having a primary circuit , the input of which is connected to means for taking off compressed air from the engine , and the output is connected to means for supplying air to the aircraft , and a secondary circuit which is supplied with air which is taken off in said air flow.2. Turbine engine according to claim 1 , wherein the exchanger is fixed to an outer annular housing of the engine.3. Turbine engine according to claim 2 , wherein the outer annular housing is configured to define the inside of said flow duct for the air flow.4. Turbine engine according to claim 2 , wherein the outer annular housing is surrounded by nacelle walls and/or cowls which define an annular space around said outer annular housing claim 2 , and wherein the heat exchanger is mounted in said annular space.5. Turbine engine according to claim 2 , wherein the outer annular housing comprises at least one recess for accommodating the exchanger which is formed by a local deformation of the housing.6. Turbine engine according to claim 5 , wherein the exchanger is fixed to a removable panel of the housing claim 5 , said panel being designed to define said recess.7. Turbine engine according to claim 2 , wherein the exchanger is embedded in part in the housing ...

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Номер записи: 34
07-01-2016 дата публикации

COMPACT AERO-THERMO MODEL STABILIZATION WITH COMPRESSIBLE FLOW FUNCTION TRANSFORM

Номер: US20160003164A1
Автор: Karpman Boris, Shishkin Serge L.
Принадлежит:

Systems and methods for controlling a fluid based engineering system are disclosed. The systems and methods may include a model processor for generating a model output, the model processor including a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode, wherein the open loop model generates a current state model as a function of the dynamic states and the model input, wherein a constraint on the current state model is based a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control system and including a series of utilities, the utilities are based on mathematical abstractions of physical properties associated with the component. The series of cycle synthesis modules may include a flow module for mapping a flow curve relating a compressible flow function to a pressure ratio and for defining a solution point located on the flow curve and a base point located off the flow curve. 1100. A control system () , comprising:{'b': 124', '130', '124, 'an actuator () for positioning a control device (), the control device defining a flow path through an aperture, the aperture defining a pressure drop along the flow path, and comprising a control surface, wherein the actuator () positions the control surface in order to regulate fluid flow across the pressure drop based on a model state;'}{'b': 111', '124, 'a control law () for directing the actuator () as a function of a model output; and'}{'b': '110', 'claim-text': [{'b': '220', 'an input object () for processing model input and setting a model operating mode;'}, {'b': 420', '110', '410, 'a set state module () for setting dynamic states of the model processor (), the dynamic states input to an open loop model () based on the model operating mode;'}, {'b': 410', '130, 'wherein the open loop model () generates a current state model as a function of the ...

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Номер записи: 35
07-01-2016 дата публикации

COMPACT AERO-THERMO MODEL BASED CONTROL SYSTEM

Номер: US20160003165A1
Автор: Dinsmore Ian Michael, Donald Matthew, Karpman Boris, Luppold Robert H., Niemczycki Thomas, Sembiante David
Принадлежит: UNITED TECHNOLOGIES CORPORATION

Systems and methods for controlling a fluid based engineering system are disclosed. The systems and methods may include a model processor for generating a model output, the model processor including a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode, wherein the open loop model generates a current state model as a function of the dynamic states and the model input, wherein a constraint on the current state model is based a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control system and including a series of utilities, the utilities are based on mathematical abstractions of physical properties associated with the component. The model processor may further include an estimate state module for determining an estimated state of the model based on a prior state model output and the current state model of the open loop model. 1. A control system , comprising:an actuator for positioning a control device comprising a control surface, wherein the actuator positions the control surface in order to control a model state;a control law for directing the actuator as a function of a model output; anda model processor for generating the model output, the model processor comprising:an input object for processing model input and setting a model operating mode;a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode;wherein the open loop model generates a current state model as a function of the dynamic states and the model input, wherein a constraint on the current state model is based on a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control device and comprising a series of utilities, the utilities based on mathematical ...

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Номер записи: 36
07-01-2016 дата публикации

GAS TURBINE ENGINE TURBINE IMPELLER PRESSURIZATION

Номер: US20160003166A1
Автор: "OConnor John J.", Chrisikos James P., Gates Roger, GLASSPOOLE David, Mosley John H., Wu Charles C.
Принадлежит:

A cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage. 1. A cooling system for a gas turbine engine turbine section , comprising:a rotor supporting a blade having a cooling passage;a disc secured relative to the rotor and forming a cavity between the rotor and the disc;a bleed air source in fluid communication with the cavity; andan impeller is arranged in the cavity and configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.2. The cooling system according to claim 1 , wherein the blade is in a last stage of a high pressure turbine section.3. The cooling system according to claim 2 , comprising a spool claim 2 , the rotor and the disc affixed to the spool for rotation therewith.4. The cooling system according to claim 1 , wherein the bleed air source is a stage of a high pressure compressor section.5. The cooling system according to claim 4 , wherein the high pressure compressor section includes an aft hub having an aft hub leak path claim 4 , the aft hub leak path in fluid communication with the cavity and configured to provide aft hub fluid to the cavity.6. The cooling system according to claim 1 , comprising a tangential on board injector having a TOBI leak path claim 1 , the TOBI leak path in fluid communication with the cavity and configured to provide a TOBI fluid to the cavity.7. The cooling system according to claim 1 , wherein the impeller is mounted on the disc.8. The cooling system according to ...

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Номер записи: 37
04-01-2018 дата публикации

SYSTEM AND METHOD FOR GAS BEARING SUPPORT OF TURBINE

Номер: US20180003080A1
Автор: Becker Thomas Lee, Ertas Bugra Han, Miller Brandon Wayne, Niergarth Daniel Alan
Принадлежит:

A bearing assembly for a turbine engine includes a first gas bearing configured to receive a load from a rotating shaft of the turbine engine, a transmission disk configured to receive the load from the first gas bearing, and a damping member coupled to a casing of a combustor section of the turbine engine. The transmission disk includes a gas delivery disk, which includes an axial opening configured to facilitate an axial flow through the gas delivery disk and a duct configured to facilitate a radial flow through the gas delivery disk to form the first gas bearing. The damping member is configured to receive the load from the transmission disk. 1. A bearing assembly for a turbine engine comprising:a first gas bearing configured to receive a load from a rotating shaft of the turbine engine; an axial opening configured to facilitate an axial flow through the gas delivery disk; and', 'a duct configured to facilitate a radial flow through the gas delivery disk to form the first gas bearing; and, 'a transmission disk configured to receive the load from the first gas bearing, wherein the transmission disk comprises a gas delivery disk, the gas delivery disk comprisesa damping member coupled to a casing of a combustor section of the turbine engine, wherein the damping member is configured to receive the load from the transmission disk.2. The bearing assembly of claim 1 , wherein the damping member is coupled to a fuel nozzle of the turbine engine.3. The bearing assembly of claim 1 , wherein the damping member is coupled to a casing of the turbine engine.4. The bearing assembly of claim 1 , wherein the damping member comprises a second gas bearing.5. The bearing assembly of claim 4 , wherein the damping member comprises a support coupled between the combustor section and the second gas bearing claim 4 , and a first stiffness of the support is less than a second stiffness of the second gas bearing.6. The bearing assembly of claim 4 , wherein a gas flow to the second gas ...

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Номер записи: 38
02-01-2020 дата публикации

METHOD OF REGULATING AIR PRESSURE IN ANTI-ICING SYSTEM

Номер: US20200003117A1
Автор: Goodman Robert, Greenberg Michael D., Zheng Zhijun
Принадлежит:

An anti-icing system of a nacelle inlet of an engine of an aircraft includes first and second direct acting valves and first and second control valve assemblies fluidly connected to the nacelle inlet. The first direct acting valve includes a first inlet, outlet, valve chamber, and piston. The first piston is positioned in the first direct acting valve. The first control valve assembly is fluidly connected to the first valve. The second direct acting valve includes a second inlet, outlet, valve chamber, and piston. The second piston is positioned in the second direct acting valve. The second direct acting valve is fluidly connected to the first direct acting valve in a series configuration. The second control valve assembly is fluidly connected to the second valve chamber. 1. A method of regulating air pressure in an anti-icing system of a nacelle inlet of an engine of an aircraft , the method comprising: a first direct acting valve with a first valve chamber, a first internal valve body, and a first piston slidably engaged with the first internal valve body;', 'a first control valve assembly with a first solenoid valve and fluidly connected to the first valve chamber of the first direct acting valve;', 'a second direct acting valve with a second valve chamber, a second internal valve body, and a second piston slidably engaged with the second internal valve body, wherein the second direct acting valve is fluidly connected to the first direct acting valve in a series configuration;', 'a second control valve assembly with a second solenoid valve and fluidly connected to the second valve chamber of the second direct acting valve; and, 'flowing air into a valve assembly comprising adjusting at least one of the first control valve assembly and the second control valve assembly in response to the temperature of the air in the outlet of the second direct acting valve by controlling an amount of electric current fed into at least one of the first solenoid valve in the first ...

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Номер записи: 39
07-01-2021 дата публикации

PARTICULATE INGESTION SENSOR FOR GAS TURBINE ENGINES

Номер: US20210003075A1
Автор: Birnkrant Michael J., Gilson Jonathan, Gorski Michael T., Lincoln David L, Schmidt Wayde R., Wang Xuemei, Wood Coy Bruce
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A multi-angle multi-wave array may comprise a first set of sensing elements, a second set of sensing elements, and a third set of sensing elements wherein the first set of sensing elements, the second set of sensing elements, and the third set of sensing elements are collectively configured to detect and discriminate between categories of foreign object debris including solid objects and particulates including silicate sand, water vapor, dust, volcanic ash, and smoke. 1. A multi-angle , multi-wave array , comprisinga first set of sensing elements, a second set of sensing elements, and a third set of sensing elements,wherein the first set of sensing elements, the second set of sensing elements, and the third set of sensing elements are collectively configured to detect and discriminate between categories of foreign object debris including solid objects and particulates including silicate sand, water vapor, dust, volcanic ash, and smoke,wherein the first set of sensing elements comprises a first light sensor, a second light sensor, and a third light sensor each aligned along a common axis,wherein the second set of sensing elements and the third set of sensing elements are arranged in at least one of circumferentially about a center defined by the second light sensor or in a ladder pattern.2. The multi-angle claim 1 , multi-wave array of claim 1 , wherein the first set of sensing elements comprises a first infrared light source and a first blue light source claim 1 ,wherein the first infrared light source and the first blue light source are each divided by the common axis and located circumferentially about the center defined by the second light sensor,{'b': 1', '2, 'wherein the first infrared light source and the first blue light source define, respectively, an angle θ and an angle θ between the common axis.'}3. The multi-angle claim 2 , multi-wave array of claim 2 , further comprising a second infrared light source and a second blue light source claim 2 , wherein the ...

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Номер записи: 40
07-01-2021 дата публикации

COMPACT AERO-THERMO MODEL BASED ENGINE POWER CONTROL

Номер: US20210003078A1
Автор: Case Thomas E., Donald Matthew, Karpman Boris, Meisner Richard P.
Принадлежит:

Systems and methods for controlling a fluid-based system are disclosed. The systems and methods may include generating a model output using a model processor, processing a model input vector and setting a model operating mode, and setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode. Synthesized parameters are generated as a function of the dynamic states and the model input vector based on a series of utilities, where at least one of the utilities is a configurable utility including one or more sub-utilities. An estimated state of the model is determined based on at least one of a prior state and the synthesized parameters. An actuator associated with the control device is directed as a function of a model output, where the model output includes an estimated thrust value for the control device. 1. A control system , comprising:an actuator operable to adjust a control device; and process a model input vector and set a model operating mode;', 'set dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode;', 'generate a plurality of synthesized parameters as a function of the dynamic states and the model input vector based on a series of utilities, wherein at least one of the utilities is a configurable utility comprising one or more sub-utilities;', 'determine an estimated state of the model based on at least one of a prior state and the synthesized parameters; and', 'process at least the synthesized parameters of the model to determine the model output comprising an estimated thrust value for the control device., 'a computer processor configured to execute a control law to control the actuator as a function of a model output and generate the model output using a model processor, wherein the model processor comprises a plurality of executable instructions to2. The control system of claim 1 , wherein the control law compares the ...

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Номер записи: 41
07-01-2021 дата публикации

HYBRID GAS TURBINE ENGINE WITH BLEED SYSTEM IMPROVEMENTS

Номер: US20210003080A1
Автор: Woods Adam
Принадлежит:

An architecture for powering systems on an aircraft has a gas turbine engine including a main compressor, a combustor, and a turbine. The turbine powers the main compressor, and further powers a propulsor. The turbine is operably connected to drive a generator. The generator is connected to store generated power at a battery. The battery is connected to provide power to a motor from the propulsor such that the propulsor can be selectively driven by both the motor and the turbine. A bleed air control system and a tap for selectively tapping compressed air from the main compressor, and a control valve for delivering at least one of the tapped compressed air or a compressed alternative air to bleed systems on an associated aircraft. An electric bleed compressor selectively compresses the compressed alternative air. The electric bleed compressor is powered by the battery. A control for controlling the control valve to selectively deliver at least one of the tapped compressed air and the compressed alternative air to the bleed systems. An aircraft is also disclosed. 1. An architecture for powering systems on an aircraft comprising:a gas turbine engine including a main compressor, a combustor, and a turbine, said turbine powering said main compressor, and further powering a propulsor;said turbine also being operably connected to drive a generator, said generator being connected to store generated power at a battery, said battery being connected to provide power to a motor to power said propulsor such that said propulsor can be selectively driven by both said motor and said turbine;a bleed air control system, and a tap for selectively tapping compressed air from said main compressor, and a control valve for delivering at least one of said tapped compressed air or a compressed alternative air to bleed systems on an associated aircraft; andan electric bleed compressor for selectively compressing the compressed alternative air, said electric bleed compressor powered by said ...

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Номер записи: 42
03-01-2019 дата публикации

INTER-TURBINE DUCTS WITH MULTIPLE SPLITTER BLADES

Номер: US20190003325A1
Автор: Couey Paul, Kuchana Vinayender, Malak Malak Fouad, Mckeever Craig, Srinivasan Balamurugan
Принадлежит: HONEYWELL INTERNATIONAL INC.

A turbine section of a gas turbine engine is provided. The turbine section includes a first turbine with a first inlet and a first outlet; second turbine with a second inlet and a second outlet; an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine, the inter-turbine duct being defined by a hub and a shroud; and at least two splitter blades disposed within the inter-turbine duct. The at least two splitter blades include a first splitter blade and a second splitter blade radially interior to the first splitter blade. At least the second splitter blade has a radial position that is greater than 60% of a distance from the shroud to the hub. 1. A turbine section of a gas turbine engine , the turbine section being annular about a longitudinal axis , the turbine section comprising:a first turbine with a first inlet and a first outlet;a second turbine with a second inlet and a second outlet;an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine, the inter-turbine duct being defined by a hub and a shroud; andat least two splitter blades disposed within the inter-turbine duct, the at least two splitter blades including a first splitter blade and a second splitter blade radially interior to the first splitter blade, wherein at least the second splitter blade has a radial position that is greater than 60% of a distance from the shroud to the hub.2. The turbine section of claim 1 , wherein the radial position of the second splitter blade is approximately 67% of the distance from the shroud to the hub.3. The turbine section of claim 2 , wherein the radial position of the first splitter blade is approximately 33% of the distance from the shroud to the hub.4. The turbine section of claim 1 , wherein the radial position of the second splitter blade is approximately 75% of the ...

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Номер записи: 43
03-01-2019 дата публикации

TURBOMACHINE COMPRISING A SURFACE AIR-OIL HEAT EXCHANGER BUILT INTO AN INTER-FLOW COMPARTMENT

Номер: US20190003390A1
Автор: ANCUTA Carmen Gina, RAMOS Bruna Manuela
Принадлежит: SAFRAN AIRCRAFT ENGINES

A turbomachine comprises an inter-flow compartment extending radially between first and second intermediate walls, gas bypass ducts passing through the second intermediate wall and communicating with evacuation outlets opening up into a secondary gas flow; an oil reservoir arranged in the inter-flow compartment, and a surface air-oil heat exchanger communicating with the oil reservoir for fluid circulation. Circumferentially in relation to the inter-flow compartment, the surface air-oil heat exchanger extends at least partially between the evacuation outlets which pass through the first intermediate wall. 1. A turbomachine comprising:a primary gas flow, and a secondary gas flow located around the primary gas flow,a first intermediate wall to radially inwardly delimit the secondary gas flow,a second intermediate wall to radially outwardly delimit the primary gas flow,an inter-flow compartment extending radially between the first and second intermediate walls, air bypass ducts of the primary gas flow:passing through the second intermediate wall and into the inter-flow compartment, andopening up into the secondary gas flow, through the first intermediate wall, via several evacuation outlets,an oil reservoir at least partially arranged in the inter-flow compartment, anda surface air-oil heat exchanger that has fluid communication with the oil reservoir and which has a heat exchange surface facing the secondary gas flow,wherein, circumferentially in relation to the inter-flow compartment, the surface air-oil heat exchanger extends at least partially between the evacuation outlets which pass through the first intermediate wall.2. The turbomachine of claim 1 , wherein claim 1 , circumferentially claim 1 , between the evacuation outlets claim 1 , the surface air-oil heat exchanger comprises several portions to achieve a heat exchange with the gas of the secondary gas flow claim 1 , said several portions being connected one to another claim 1 , at least two by two claim 1 , ...

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Номер записи: 44
03-01-2019 дата публикации

GAS TURBINE COOLING SYSTEM, GAS TURBINE FACILITY INCLUDING THE SAME, AND CONTROL METHOD OF GAS TURBINE COOLING SYSTEM

Номер: US20190003394A1
Автор: Fujii Satoko, Higashi Kazuya, YABE Tetsuya
Принадлежит:

A gas turbine cooling system includes: a cooling air line that guides compressed air compressed by an air compressor to a hot part; a cooler that cools the compressed air in the cooling air line; a return line that returns cooling air in the cooling air line to an upstream side in the cooling air line; a return valve that adjusts the flow rate of the cooling air flowing through the return line; and a control device that controls the degree of opening of the return valve. The control device has a second valve command generation section that, when a reception unit receives a load rejection command, generates as a second valve command a valve command ordering the degree of opening of the return valve to be forcedly increased to a predetermined load rejection-adapted degree of opening. 1. A gas turbine cooling system comprising:a cooling air line that guides compressed air compressed by an air compressor of a gas turbine to a hot part coming in contact with combustion gas in the gas turbine;a cooler that cools the compressed air in the cooling air line to produce cooling air;a booster that pressurizes the cooling air in the cooling air line;a return line that returns the cooling air in a discharge line that is a line of the cooling air line located on a side of the hot part from the booster, to an intake air line that is a line of the cooling air line located on a side of the air compressor from the booster;a return valve that is provided in the return line and adjusts a flow rate of the cooling air flowing through the return line;a detector that detects a state amount of the cooling air flowing through the intake air line and a state amount of the cooling air flowing through the discharge line; and a reception unit that receives a load rejection command indicating a load rejection of the gas turbine;', 'a first valve command generation section that generates a first valve command indicating a degree of opening of the return valve according to the state amount detected ...

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Номер записи: 45
03-01-2019 дата публикации

GAS TURBINE ENGINE VARIABLE AREA FAN NOZZLE CONTROL

Номер: US20190003422A1
Автор: Blackwell Geoffrey T., McVey Wililam J., Tempelman William G.
Принадлежит:

A method of managing a gas turbine engine includes the steps of detecting an airspeed and detecting a fan speed. A parameter relationship is referenced related to a desired variable area fan nozzle position based upon at least airspeed and fan speed. The detected airspeed and detected fan speed is compared to the parameter relationship to determine a target variable area fan nozzle position. An actual variable area fan nozzle position is adjusted in response to the determination of the target area fan nozzle position and at least one threshold. 1. A method of managing a gas turbine engine comprising the steps of:detecting an airspeed;detecting a fan speed;referencing a parameter relationship related to a desired variable area fan nozzle position based upon at least airspeed and fan speed, and comparing the detected airspeed and detected fan speed to the parameter relationship to determine a target variable area fan nozzle position; andadjusting an actual variable area fan nozzle position in response to the determination of the target area fan nozzle position and at least one threshold.2. The method according to claim 1 , wherein the fan speed detecting step includes detecting a low speed spool rotational speed claim 1 , and correcting the fan speed based upon an ambient temperature.3. The method according to claim 2 , wherein the fan speed detecting step includes calculating the fan speed based upon a gear reduction ratio.4. The method according to claim 1 , wherein the referencing and comparing steps include providing a target variable area fan nozzle position for a range of air speeds based upon the fan speed.5. The method according to claim 4 , wherein the air speed range is 0.35-0.55 Mach claim 4 , and the data table includes first and second thresholds corresponding to lower and upper fan speed limits claim 4 , the target variable area fan nozzle position selected based upon the first and second thresholds.6. The method according to claim 5 , wherein the upper ...

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Номер записи: 46
07-01-2021 дата публикации

METHOD AND SYSTEM FOR MEASURING TEMPERATURE IN A GAS TURBINE ENGINE

Номер: US20210003458A1
Автор: BOUCHARD Guy, DUROCHER ERIC, Smith Scott
Принадлежит:

A system and method for measuring average temperature of gas in an axial cross-section of a gas turbine engine gas path, involving diverting gas samples from different positions in the axial cross-section to a gas mixing chamber and measuring a temperature of the resulting mixed gas. 1. A method for measuring average temperature of gas in an axial cross-section of a gas turbine engine gas path comprising:diverting a plurality of gas samples from the gas path from positions in the axial cross-section circumferentially spaced from one another;mixing the gas samples to form a mixed gas; andmeasuring a temperature of the mixed gas.2. The method of wherein the corresponding positions lie in a common plane.3. The method of wherein the corresponding positions are equally interspaced from one another around the axial cross-section.4. The method of claim 1 , claim 1 , or further comprising conveying the diverted gas samples to a gas mixing chamber.5. The method of claim 1 , claim 1 , or further comprising conveying the mixed gas to a gas measuring chamber claim 1 , wherein the measuring is performed in the gas measuring chamber.6. The method of claim 1 , claim 1 , or further comprising releasing the mixed gas to the gas path after said measuring.7. The method of further comprising releasing the mixed gas upstream of a low-pressure turbine rotor claim 6 , after said measuring.8. The method of further comprising releasing the mixed gas downstream of a low-pressure turbine rotor claim 6 , after said measuring.9. The method of claim 6 , claim 6 , or further comprising releasing the mixed gas into a space external to the gas turbine engine claim 6 , after said measuring.10. The method of claim 6 , claim 6 , or wherein the axial cross-section is in a turbine section of the turbine engine.11. A gas turbine engine having a gas path extending sequentially across a compressor section claim 6 , a combustor claim 6 , and a turbine section claim 6 , the gas path being annular around an ...

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Номер записи: 47
20-01-2022 дата публикации

DEVICES AND METHODS FOR GUIDING BLEED AIR IN A TURBOFAN ENGINE

Номер: US20220018292A1
Автор: ANAND Karan, CAPRON Alexandre, YAM Tammy
Принадлежит:

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 ...

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Номер записи: 48
20-01-2022 дата публикации

Compressor apparatus with bleed slot and supplemental flange

Номер: US20220018293A1
Автор: Atanu Saha, Manish Singhal, Michael Anthony Thomas, Steven Mitchell Taylor, Vishnu Das K
Принадлежит: General Electric Co

An example compressor bleed slot apparatus includes an annular compressor casing, a blade row mounted for rotation about a centerline axis inside the compressor casing, a bleed slot passing through the forward section of the compressor casing, wherein the bleed slot is bounded by inboard and outboard walls defined within the compressor casing, the bleed slot extending along a slot axis, at least a portion of the bleed slot lying within an axial extent of the blade row, an array of struts interconnecting the inboard and outboard walls, and an annular supplemental flange extending radially outward from the forward section of the compressor casing, wherein at least a portion of the supplemental flange is axially positioned within an axial extent of the bleed slot, wherein the supplemental flange includes a necked-down portion adjacent the forward section of the compressor casing, the necked-down portion positioned axially between an inlet of the bleed slot and an outlet of the bleed slot.

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Номер записи: 49
12-01-2017 дата публикации

Supply of air to an air-conditioning circuit of an aircraft cabin from its turboprop engine

Номер: US20170008633A1
Автор: Bruno Thoraval, Jean-Francois Rideau, Nicolas Claude PARMENTIER, Pierre Froment
Принадлежит: Safran Aircraft Engines SAS, Safran Power Units SAS, Safran SA

An aircraft turboprop engine includes at least a low-pressure body and a high-pressure body. The low-pressure body drives a propeller by means of a gearbox. The turboprop engine also includes means for supplying air to an air-conditioning circuit of an aircraft cabin, wherein the supply means has at least one compressor borne by the gearbox and of which the rotor is coupled to the low-pressure body by means of the gearbox.

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Номер записи: 50
12-01-2017 дата публикации

INTEGRATION OF PRESSURE SWING ADSORPTION WITH A POWER PLANT FOR CO2 CAPTURE/UTILIZATION AND N2 PRODUCTION

Номер: US20170009652A1
Автор: Caram Hugo S., GUPTA Ramesh, Mittricker Franklin F., Starcher Loren K., SUNDARAM Narasimhan, THOMANN Hans, Weigel Scott J., Weston Simon C.
Принадлежит:

Systems and methods are provided for combined cycle power generation while reducing or mitigating emissions during power generation. Recycled exhaust gas from a power generation combustion reaction can be separated using a swing adsorption process so as to generate a high purity COstream while reducing/minimizing the energy required for the separation and without having to reduce the temperature of the exhaust gas. This can allow for improved energy recovery while also generating high purity streams of carbon dioxide and nitrogen. 119.-. (canceled)20. A method for production of Nand COfrom a reactor exhaust stream , comprising:{'sub': 2', '2, 'passing a reactor exhaust stream comprising at least about 70 vol % Nand at least about 10 vol % COinto a swing adsorption reactor comprising an adsorbent material, the reactor exhaust stream having a pressure between about 10 bara (about 1.0 MPaa) to about 30 bara (about 3.0 MPaa);'}{'sub': '2', 'adsorbing COon the adsorbent material at an adsorption temperature of at least 400° C.;'}{'sub': 2', '2, 'recovering an Nstream with a purity of at least about 95 vol % from a forward end of the reactor, the recovered Nstream having a pressure that differs from the pressure of the reactor exhaust stream by about 0.5 bar (about 50 kPa) or less;'}reducing the pressure in the swing adsorption reactor to a pressure from about 1.0 bara (about 0.1 MPaa) to about 4.0 bara (about 0.4 MPaa) by outputting a blow down stream from at least one end of the reactor; and{'sub': 2', '2', '2', '2', '2, 'purging the swing adsorption reactor with a steam purge at a pressure from about 1.0 bara (about 0.1 MPaa) to about 4.0 bara (about 0.4 MPaa) to generate a COrecovery stream, the COrecovery stream comprising at least about 90% of the COpresent in the reactor exhaust stream, the steam purge containing less than about 1.0 moles of HO per mole of COin the reactor exhaust stream.'}21. The method of claim 20 , wherein the passing claim 20 , the adsorbing ...

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Номер записи: 51
12-01-2017 дата публикации

COMPRESSOR ENDWALL BOUNDARY LAYER REMOVAL

Номер: US20170009663A1
Автор: Epstein Alan H.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A compressor is provided. The compressor may comprise a blade configured to rotate about an axis, an inner endwall coupled to the blade, and an outer endwall radially outward from the blade and the inner endwall. The outer endwall and the inner endwall may define a flow path. A vane may be disposed aft of the blade and coupled to the outer endwall, and a bleed passage may be selectively positioned to extract a boundary layer flow. A method of locating bleed passages is also provided. The method may include comprise analyzing at least one an entropy or a temperature of a boundary layer flow, identifying endwall locations where the entropy or the temperature are elevated, and forming a bleed passage at the endwall locations. 1. A compressor , comprising:a blade configured to rotate about an axis;an inner endwall coupled to the blade;an outer endwall radially outward from the blade and the inner endwall, the outer endwall and the inner endwall defining a flow path;a vane aft of the blade and coupled to the outer endwall;a bleed passage selectively positioned on at least one of the inner endwall, the outer endwall, or the vane to extract a boundary layer flow.2. The compressor of claim 1 , wherein the bleed passage extends through the inner endwall aft of the blade.3. The compressor of claim 2 , wherein the bleed passage comprises an cylindrical geometry.4. The compressor of claim 2 , wherein the bleed passage comprises an opening flush with the inner endwall.5. The compressor of claim 1 , wherein the bleed passage extends through the outer endwall aft of the blade.6. The compressor of claim 5 , wherein the bleed passage comprises an opening flush with the outer endwall.7. The compressor of claim 1 , wherein the bleed passage is disposed on the vane.8. The compressor of claim 7 , wherein the vane is cantilevered from the outer endwall.9. The compressor of claim 7 , further comprising a passage inside the vane and in fluid communication with the bleed passage.10. A gas ...

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Номер записи: 52
11-01-2018 дата публикации

Bleed flow extraction system for a gas turbine engine

Номер: US20180009536A1
Автор: Brandon Wayne Miller, Charles Kammer Christopherson
Принадлежит: General Electric Co

An air cycle machine for extracting bleed air from a gas turbine engine of an aircraft is provided. The air cycle machine extracts a stream of low pressure bleed air and a stream of high pressure bleed air from a compressor section of the gas turbine engine. The air cycle machine includes a compressor that receives the stream of low pressure bleed air and a turbine that receives the stream of high pressure bleed air. The stream of high pressure bleed air is expanded as it drives the turbine, and the stream of low pressure bleed air is compressed by the compressor. The resulting streams of bleed air are substantially the same pressure, such that they may be merged using a junction into a combined bleed air stream having a temperature and pressure suitable for use by a variety of aircraft accessory systems, such as an environmental control system. The air cycle machine may further power or be powered from an electrical storage device or generator on the fan.

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Номер записи: 53
27-01-2022 дата публикации

SYSTEM AND METHODS FOR CONTROLLING SURGE MARGIN IN THE COMPRESSOR SECTION OF A GAS TURBINE ENGINE

Номер: US20220025824A1
Автор: Burkholder Graham Adam, Iwrey Benjamin Michael
Принадлежит:

Systems and methods are disclosed for controlling surge margin in the compressor section of a gas turbine engine. A first compressor section and a second compressor section are in fluid communication with a bypass conduit. An auxiliary turbine and discharge conduit are positioned in the bypass conduit. Fluid flow from the compressor sections into the bypass conduit is controlled by bypass control valves. 1. A system for controlling surge margin in a compressor section of a gas turbine engine , the system comprising: one or more compressor stages defining a first compressor section flowpath; and', 'a first compressor section discharge in fluid communication with the first compressor section flowpath;, 'a first compressor section comprisinga first compressor section bypass port positioned along the first compressor section and in fluid communication with the first compressor flowpath; one or more compressor stages defining a second compressor section flowpath, the second compressor section flowpath in fluid communication with the first compressor section flowpath; and', 'a second compressor section discharge in fluid communication with the second compressor section flowpath;, 'a second compressor section comprisinga second compressor section bypass port positioned along the second compressor section and in fluid communication with the second compressor section flowpath;a bypass conduit, extending between the first compressor section bypass port and the second compressor section bypass port;a first compressor section bypass control valve positioned in the bypass conduit downstream of the first compressor section bypass port;a second compressor section bypass control valve positioned in the bypass conduit downstream of the second compressor section bypass port;an auxiliary turbine positioned in the bypass conduit between the first compressor section bypass control valve and the second compressor section bypass control valve;a discharge conduit coupled to and in fluid ...

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Номер записи: 54
14-01-2016 дата публикации

VENTILATION INLET

Номер: US20160010558A1
Автор: HUSSAIN Zahid M., IRELAND Peter T., MISHRA Nityabh
Принадлежит:

A ventilation inlet including a conduit () arranged to convey flow from a first flow zone to a second flow zone. The conduit has a mouth region () presenting to the first flow zone an entrance aperture to receive the flow therefrom. The conduit has a baffle () spanning a portion of the conduit to define a throat region (), the throat region being narrower than the entrance aperture. The throat region () is movable along the conduit () to control the flow through the ventilation inlet. 1. A ventilation inlet including a conduit arranged to convey flow from a first flow zone to a second flow zone , the conduit having a mouth region presenting to the first flow zone an entrance aperture to receive the flow therefrom; anda baffle spanning a portion of the conduit to define a throat region, the throat region being narrower than the entrance aperture; whereinthe throat region is movable along the conduit to control the flow through the ventilation inlet.2. A ventilation inlet according to wherein the baffle is movable along the conduit.3. A ventilation inlet according to wherein the baffle is movable along the general direction of flow in the conduit.4. A ventilation inlet according to wherein the baffle spans transversely across a portion of the conduit.5. A ventilation inlet according to wherein the baffle is locatable in a first position claim 1 , in which the baffle is located downstream of the mouth region so that the throat region is recessed from the entrance aperture.6. A ventilation inlet according to wherein the baffle is movable to a second position claim 1 , in which the baffle is located in the mouth region so that the throat region is presented to the first flow zone to receive the flow instead of the entrance aperture of the mouth region.7. A ventilation inlet according to wherein the conduit has a central axis extending generally in the direction of flow through the conduit.8. A ventilation inlet according to wherein the baffle is arranged to extend at ...

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Номер записи: 55
14-01-2016 дата публикации

BLEED VALVE ASSEMBLY

Номер: US20160010564A1
Автор: LIOR Raphael, PAVLIK Mark G.
Принадлежит:

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: ...

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Номер записи: 56
14-01-2016 дата публикации

GAS TURBINE ENGINE WITH FAN VARIABLE AREA NOZZLE FOR LOW FAN PRESSURE RATIO

Номер: US20160010565A9
Автор: Kohlenberg Gregory A., Schwarz Frederick M., Zamora Sean P.
Принадлежит:

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45. 1. A gas turbine engine comprising:a core nacelle defined about an engine centerline axis;a core engine at least partially disposed within the core nacelle;a fan section with twenty (20) or less fan blades;a gear system driven by the core engine to drive said fan section;a fan nacelle mounted at least partially around said fan section and said core nacelle to define a fan bypass flow path for a fan bypass airflow, said fan bypass airflow having a fan pressure ratio of the fan bypass airflow during engine operation, said fan pressure ratio less than about 1.45;a variable fan nozzle axially movable relative to the fan nacelle, the variable fan nozzle including at least two sectors; anda controller for independently adjusting each of the at least two sectors.2. (canceled)3. The engine as recited in claim 1 , wherein the controller is operable to reduce said fan nozzle exit area at a cruise flight condition.4. The engine as recited in claim 1 , wherein said controller is operable to control said fan nozzle exit area to reduce a fan instability.5. The engine as recited in claim 1 , wherein said fan variable area nozzle defines a trailing edge of said fan nacelle.6. The engine as recited in claim 1 , wherein said fan variable area nozzle is axially movable relative to said fan nacelle.7. (canceled)8. The engine as recited in claim 1 , wherein said fan section defines a corrected fan tip speed less than about 1150 ft/second.9. The engine as recited in claim 1 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than about five (5).10. The engine as recited in claim 7 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than five (5).11. The engine as recited in claim 1 , further comprising a gear system ...

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Номер записи: 57
09-01-2020 дата публикации

METHOD AND DEVICE FOR DETECTING CONDITIONS CONDUCIVE TO THE ONSET OF PUMPING WITH A VIEW TO PROTECTING A COMPRESSOR OF AN AIRCRAFT TURBINE ENGINE

Номер: US20200010211A1
Автор: Eburderie Emmanuel Mickaël
Принадлежит: SAFRAN AIRCRAFT ENGINES

A method and device for detecting conditions conducive to the onset of pumping that can affect a low-pressure compressor of an aircraft turbine engine. The turbine engine including a high-pressure compressor. The method including measuring a speed variation of the aircraft and measuring a speed variation of the high-pressure compressor. The method including a preliminary step of measuring an altitude of the aircraft. The conditions conducive to the onset of pumping being detected when the following conditions are jointly obtained: (a) the speed variation measured over a predetermined time interval corresponds to an acceleration greater than a first positive threshold, (b) the measured speed variation corresponds to a deceleration less than a second negative threshold, and (c) the altitude is greater than a third predetermined threshold. 1. A method for detecting conditions conducive to an onset of pumping affecting a low-pressure compressor of a turbine engine for an aircraft , said turbine engine further comprising a high-pressure compressor , the method comprising:measuring a speed variation of said aircraft;measuring a speed variation of said high-pressure compressor;measuring an altitude of said aircraft;determining the conditions conducive to the onset of pumping when (a) said speed variation measured over a predetermined time interval corresponds to an acceleration greater than a first positive threshold, (b) said measured speed variation corresponds to a deceleration less than a second negative threshold; and (c) said measured altitude is higher than a predetermined third threshold.2. The method according to claim 1 , wherein said first threshold is between 10Mach per second and 10Mach per second.3. The method according to claim 1 , wherein said time interval is between 3 seconds and 20 seconds.4. The method according claim 1 , wherein said second threshold is between −2 revolutions per minute per second and −20 revolutions per minute per second.5. The method ...

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Номер записи: 58
11-01-2018 дата публикации

NACELLE ANTI ICE SYSTEM

Номер: US20180010519A1
Автор: Goodman Robert, Greenberg Michael D., Zheng Zhijun
Принадлежит:

An anti-icing system of a nacelle inlet of an engine of an aircraft includes first and second direct acting valves and first and second control valve assemblies fluidly connected to the nacelle inlet. The first direct acting valve includes a first inlet, outlet, valve chamber, and piston. The first piston is positioned in the first direct acting valve. The first control valve assembly is fluidly connected to the first valve. The second direct acting valve includes a second inlet, outlet, valve chamber, and piston. The second piston is positioned in the second direct acting valve. The second direct acting valve is fluidly connected to the first direct acting valve in a series configuration. The second control valve assembly is fluidly connected to the second valve chamber. 1. An anti-icing system of a nacelle inlet of an engine of an aircraft , wherein the anti-icing system comprises: [ a first inlet;', 'a first valve chamber fluidly connected to the first inlet;', 'a first internal valve body circumferentially surrounding the first valve chamber;', 'a first outlet; and', 'a first piston for adjusting a rate of flow of air through the first direct acting valve, wherein the first piston is slidably engaged with the first internal valve body;, 'a first direct acting valve comprising, 'a first control valve assembly fluidly connected to the first valve chamber of the first direct acting valve;', a second inlet;', 'a second valve chamber fluidly connected to the second inlet;', 'a second internal valve body circumferentially surrounding the second valve chamber;', 'a second outlet; and', 'a second piston for adjusting a rate of flow of air through the second direct acting valve, wherein the second piston is slidably engaged with the second internal valve body, and, 'a second direct acting valve comprising], 'a valve assembly fluidly connected to the nacelle inlet, wherein the valve assembly comprises 'a second control valve assembly fluidly connected to the second valve ...

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Номер записи: 59
11-01-2018 дата публикации

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

Номер: US20180010520A1
Автор: AI Toshishige, FUKUNAGA Yuya, Iwasaki Tatsuya, Iwasaki Yoshifumi
Принадлежит:

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 ...

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Номер записи: 60
10-01-2019 дата публикации

AIR GUIDING SYSTEM IN AN AIRCRAFT TURBO ENGINE

Номер: US20190010817A1
Автор: KAZAKOS Charilaos
Принадлежит:

An air guiding system in an aircraft turbo engine, wherein at least one duct, in particular a pipe, guides an air flow from an offtake in the aircraft turbo engine to at least one exit in a region upstream from at least one air flow opening to at least one cavity, the at least one cavity being in flow connection with a turbine stage, in particular a low pressure turbine stage, the at least one flow opening being coupled with at least one air flow guiding element for guiding a part of the air flow through the at least one air flow opening into the at least one cavity. 1. An air guiding system in an aircraft turbo engine , wherein at least one duct , in particular a pipe , guides an air flow from an offtake in the aircraft turbo engine to at least one exit in a region upstream from at least one air flow opening to at least one cavity ,the at least one cavity is in flow connection with a turbine stage, in particular a low pressure turbine stage,the at least one flow opening being coupled with at least one air flow guiding element for guiding a part of the air flow through the at least one air flow opening into the at least one cavity.2. The air guiding system according to claim 1 , with at least one air offtake at a compressor stage in the aircraft turbo engine.3. The air guiding system according to claim 1 , wherein the distance of the at least one air duct exit to the leading edge of the at least one flow guiding element is minimized.4. The aiir guiding system according to claim 1 , wherein the at least one air duct exit having a distance to leading edge of the at least one flow guiding element between 0.1 to 5 times the characteristic diameter of the air duct exit claim 1 , in particular 0.5 to 3 times the characteristic diameter of the air duct exit.5. The air guiding system according to claim 1 , wherein the flow characteristic width of the at least one air flow guiding element is at least equal to the flow characteristic diameter of the at least one air flow ...

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Номер записи: 61
14-01-2021 дата публикации

Oxidation activated cooling flow

Номер: US20210010423A1
Автор: Lyndsay Marie Kibler, Matthew Troy HAFNER, Zachary John Snider
Принадлежит: General Electric Co

A flow regulating system for increasing a flow of cooling fluid supplied to a cooling system of a component of a gas turbine system is provided. The flow regulating system includes: a pneumatic circuit embedded within a section of the component, the pneumatic circuit including a set of interconnected pneumatic passages; and a pressure-actuated switch fluidly coupled to the pneumatic circuit. The pressure-actuated switch is activated in response to a formation of a breach in the section of the component and an exposure of at least one of the pneumatic passages of the pneumatic circuit embedded in the section of the component. The activation of the pressure-actuated switch increases the flow of cooling fluid supplied to the cooling system of the component.

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Номер записи: 62
14-01-2021 дата публикации

Takeoff power boost

Номер: US20210010436A1
Автор: Brett Partington, William Gons
Принадлежит: Bell Textron Inc

Embodiments are directed to boosting aircraft engine performance for takeoff and critical mission segments by reducing airflow used for cooling exhaust gases. The airflow is reduced by stopping an accessory blower or by closing an external air vent. Eliminating the cooling airflow to the exhaust has the effect of lowering the backpressure on the engine, which thereby increases maximum engine power.

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Номер записи: 63
10-01-2019 дата публикации

SYSTEMS AND METHODS FOR COOLING COMPONENTS OF A GAS TURBINE

Номер: US20190010869A1
Автор: Kerth Jason M.
Принадлежит:

Systems and methods for cooling one or more components of a gas turbine are provided. One system may include an expansion device and one or more conduits. The expansion device may be operatively coupled to the gas turbine and configured to convert a pressure drop of a stream of compressed process fluid to mechanical energy. The expansion device may be further configured to at least partially drive the gas turbine with the mechanical energy. The one or more conduits may fluidly couple the expansion device and the gas turbine. The one or more conduits may be configured to direct an expanded stream of the compressed process fluid to the one or more components of the gas turbine to cool the one or more components. 1. A gas turbine assembly , comprising: a combustor configured to receive a first stream of a compressed process fluid, mix a fuel with the first stream of the compressed process fluid to form a mixture, and combust the mixture to form a combustion product, and', 'a power turbine configured to receive and expand the combustion product to convert a pressure drop of the combustion product to mechanical energy;, 'a gas turbine comprising'}an expansion device configured to receive and expand a second stream of the compressed process fluid to convert a pressure drop of the second stream of the compressed process fluid to mechanical energy, the expansion device operatively coupled to the power turbine and configured to at least partially drive the power turbine with the mechanical energy converted from the pressure drop of the second stream of the compressed process fluid; anda plurality of conduits fluidly coupling the power turbine and the expansion device and configured to direct the second stream of the compressed process fluid expanded in the expansion device to the power turbine to cool one or more components of the power turbine.2. The gas turbine assembly of claim 1 , wherein the power turbine comprises a plurality of stages claim 1 , each stage comprising a ...

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Номер записи: 64
10-01-2019 дата публикации

TURBINE ENGINE THERMAL MANAGEMENT

Номер: US20190010870A1
Автор: Glover Jeffrey, Miller Brandon Wayne
Принадлежит:

A gas turbine engine including core engine is provided. Air may enter the core engine through an inlet and travel through and engine air flowpath extending through the core engine, e.g., generally along an axial direction of the gas turbine engine. The gas turbine engine additionally includes a cooling air flowpath extending outwardly generally along the radial direction of the gas turbine engine. The cooling air flowpath extends between an inlet in flow communication with engine air flowpath and an outlet defined by an opening in an outer casing of the core engine. Moreover, the gas turbine engine includes a heat exchanger positioned at least partially within the outer casing the core engine with the cooling air flowpath extending over or through the heat exchanger. 1. A gas turbine engine defining a radial direction , the gas turbine engine comprising:a core engine including an outer casing;an engine air flowpath extending through the core engine;a cooling air flowpath extending between an inlet in flow communication with the engine air flowpath and an outlet defined by an opening in the outer casing of the core engine; anda heat exchanger positioned in thermal communication with the cooling air flowpath.2. The gas turbine of claim 1 , wherein the core engine includes a vent over the opening in the outer casing claim 1 , the vent configured to adjust an amount of airflow allowable through the cooling air flowpath.3. The gas turbine of claim 1 , wherein the heat exchanger is rigidly attached to the outer casing and is configured as an air cooled oil cooler.4. The gas turbine of claim 3 , wherein the core engine includes an annular compressor frame positioned within the outer casing claim 3 , and wherein the heat exchanger is also rigidly attached to the annular compressor frame such that the heat exchanger provides structural support between the outer casing and the annular compressor frame.5. The gas turbine of claim 1 , wherein the core engine includes a ...

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Номер записи: 65
18-01-2018 дата публикации

BLEED AIR HEAT EJECTORS

Номер: US20180016017A1
Автор: daSilva Jose A., JR. Donald William, Lamb, Simonetti Joseph Lawrence, Tran Nhan R.
Принадлежит:

A bleed air heat ejector includes a housing defining an inner passage extending in an axial direction from an inlet of the housing to an outlet of the housing, wherein the inlet is configured to channel ambient air into the housing, and wherein the outlet is configured to channel mixed heated and ambient air out of the housing. A dispenser is mounted in the housing between the inlet and the outlet. The dispenser includes an inner chamber configured to receive heated air from a bleed line. The dispenser includes a plurality of apertures therein for issuing heated air from the inner chamber into the inner passage of the housing to form a flow of mixed ambient and bleed air. The apertures of the inner chamber can be angled primary jets, vortex generators can be included, and/or secondary jets can be included to promote mixing of bleed and ambient air. 1. A bleed air heat ejector comprising:a housing defining an inner passage extending in an axial direction from an inlet of the housing to an outlet of the housing, wherein the inlet is configured to channel ambient air into the housing, and wherein the outlet is configured to channel mixed heated and ambient air out of the housing;a dispenser mounted in the housing between the inlet and the outlet, wherein the dispenser includes an inner chamber configured to receive heated air from a bleed line, wherein the dispenser includes a plurality of apertures therein for issuing heated air from the inner chamber into the inner passage of the housing to form a flow of mixed ambient and bleed air; anda plurality of vortex generators mounted in the inner passage downstream of the apertures in the inner chamber, wherein the vortex generators are configured to promote mixing of bleed and ambient air.2. A bleed air heat ejector as recited in claim 1 , wherein the inner passage defines a venturi with a throat claim 1 , wherein the apertures of the inner chamber are directed downstream from the throat of the venturi claim 1 , wherein ...

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Номер записи: 66
21-01-2016 дата публикации

Compact Aero-Thermo Model Based Degraded Mode Control

Номер: US20160017813A1
Автор: Karpman Boris, Meisner Richard P.
Принадлежит:

Systems and methods for controlling a fluid based engineering system are disclosed. The systems and methods may include a model processor for generating a model output, the model processor including a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode. The system may include a control law for directing the actuator as a function of a model output and for determining if the control device is operating with deteriorated conditions. The model processor may further include an estimate state module for determining an estimated state of the model based on a prior state model output and the current state model of an open loop model. 1. A control system , comprising:an actuator for positioning a control device comprising a control surface, wherein the actuator positions the control surface in order to control a model state;a control law for directing the actuator as a function of a model output and for determining if the control device is operating with deteriorated conditions; and an input object for processing model input and setting a model operating mode;', 'a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode;', 'wherein the open loop model generates a current state model as a function of the dynamic states and the model input, a constraint on the current state model being based on a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control device and comprising a series of utilities, the utilities based on mathematical abstractions of physical properties associated with the component;', 'an estimate state module for determining an estimated state of the model based on a prior state model output and the current state model of the open loop model; and', 'an output object for processing the estimated ...

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Номер записи: 67
21-01-2016 дата публикации

Compact Aero-Thermo Model Based Engine Material Temperature Control

Номер: US20160017814A1
Автор: Dinsmore Ian Michael, Karpman Boris, Meisner Richard P., Shade John L.
Принадлежит:

Systems and methods for controlling a fluid based engineering system are disclosed. The systems and methods may include a model processor for generating a model output, the model processor including a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode, wherein the open loop model generates a current state model as a function of the dynamic states and the model input, wherein a constraint on the current state model is based a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control system and including a series of utilities, the utilities are based on mathematical abstractions of physical properties associated with the component. The series of utilities may include a material temperature utility for determining a material temperature associated with a component of the cycle of the control system The model processor may further include an estimate state module for determining an estimated state of the model based on a prior state model output and the current state model of the open loop model. 1. A control system , comprising:an actuator for positioning a control device comprising a control surface, wherein the actuator positions the control surface in order to control a model state;a control law for directing the actuator as a function of a model output; and an input object for processing model input and setting a model operating mode;', 'a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode;', 'wherein the open loop model generates a current state model as a function of the dynamic states and the model input, a constraint on the current state model being based on a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control ...

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Номер записи: 68
21-01-2016 дата публикации

EXPANDING SHELL FLOW CONTROL DEVICE

Номер: US20160017815A1
Автор: Izquierdo Felix, McAlice Timothy J.
Принадлежит:

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.

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Номер записи: 69
18-01-2018 дата публикации

RAMBURNING ENGINE WITH INLET TURBINE

Номер: US20180017017A1
Автор: Oechsle Victor Lewis, Tomlinson Steven Wesley
Принадлежит: Rolls Royce North American Technologies Inc.

A gas turbine engine assembly for high Mach number operation comprising an inlet turbine assembly, a gas turbine engine core, and a ramburner arranged along a common axis. During high Mach number operation, working fluid flow is selectively directed through the inlet turbine assembly to cool the working fluid prior to entry in the gas turbine engine core. Working fluid exiting the gas turbine engine core may be reheated by a ramburner for thrust augmentation. 1. A high Mach number engine comprising:a gas turbine core including a compressor, a combustor, and a turbine;an inlet assembly coupled to the gas turbine core and configured to selectively remove energy from a working fluid prior to entering the gas turbine core when the high Mach number engine is travelling at high speeds; anda ramburner assembly coupled to the gas turbine core and configured to selectively mix the working fluid with a bypass air upon exiting the gas turbine core and add energy to the working fluid for thrust augmentation.2. The high Mach number engine of wherein the inlet assembly comprises an inlet turbine disposed in an inlet turbine casing and an inlet flow director configured to selectively permit working fluid to flow through the inlet turbine.3. The high Mach number engine of further comprising an inlet turbine load component coupled to the inlet turbine.4. The high Mach number engine of further comprising an engine casing which encases the inlet turbine assembly and the gas turbine core claim 2 , wherein an inlet turbine passageway is defined within the inlet turbine casing and an inlet bypass passageway is defined between the engine casing and the inlet turbine casing and is coaxial with the inlet turbine casing claim 2 , and wherein the high Mach number engine further comprises a core flow director movable from (i) an open position arranged to allow working fluid from the inlet bypass passageway to move through the inlet bypass passageway and enter the gas turbine core without ...

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Номер записи: 70
21-01-2021 дата публикации

Modulated combustor bypass for hybrid idle

Номер: US20210016888A1
Автор: Neil Terwilliger, Victor Pascu
Принадлежит: Raytheon Technologies Corp

A hybrid propulsion system includes a gas turbine engine having a low speed spool, a high speed spool, and a combustor. The low speed spool includes a low pressure compressor and a low pressure turbine, and the high speed spool includes a high pressure compressor and a high pressure turbine. The hybrid propulsion system also includes a motor configured to augment rotational power of the high speed spool, a flow modulation device configured to control a combustor bypass air flow around the combustor to the turbine section, and a controller. The controller is operable to determine a mode of operation, apply supplemental power to the high speed spool using the motor, modulate the combustor bypass air flow using the flow modulation device, and adjust a fuel-air ratio at the combustor based on modulation of the combustor bypass air flow and the supplemental power applied to the high speed spool.

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Номер записи: 71
17-01-2019 дата публикации

CONTINUOUS DETONATION GAS TURBINE ENGINE

Номер: US20190017437A1
Автор: Glaser Aaron, Joshi Narendra Digamber, Pal Sibtosh, Tangirala Venkat Eswarlu
Принадлежит:

A gas turbine engine includes a primary combustor, a secondary combustor, a high pressure (HP) turbine, and a mixing duct. The HP turbine is downstream of the primary combustor and fluidly connected to a rear end of the primary combustor via a first exhaust duct. The mixing duct is disposed downstream of the HP turbine and the secondary combustor. The mixing duct has a first inlet fluidly connected to the HP turbine via a turbine exit duct, a second inlet fluidly connected to a rear end of the secondary combustor via a second exhaust duct, and an outlet. The turbine exit duct directs a primary exhaust stream, which is emitted from the primary combustor and expanded through the HP turbine, into the mixing duct. The second exhaust duct directs a secondary exhaust stream emitted from the secondary combustor into the mixing duct. 1. A gas turbine engine comprising:a primary combustor including an annular combustion chamber extending between front and rear ends of the primary combustor;a high pressure (HP) turbine downstream of the primary combustor and fluidly connected to the rear end of the primary combustor via a first exhaust duct, the first exhaust duct positioned to direct a primary exhaust stream emitted from the primary combustor to the HP turbine;a secondary combustor including an annular combustion chamber extending between front and rear ends of the secondary combustor; anda mixing duct disposed downstream of the HP turbine and the secondary combustor, the mixing duct having a first inlet fluidly connected to the HP turbine via a turbine exit duct, a second inlet fluidly connected to the rear end of the secondary combustor via a second exhaust duct, and an outlet, wherein the turbine exit duct directs the primary exhaust stream from the HP turbine into the mixing duct and the second exhaust duct directs a secondary exhaust stream emitted from the secondary combustor into the mixing duct.2. The gas turbine engine of claim 1 , wherein the secondary combustor is ...

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Номер записи: 72
17-01-2019 дата публикации

TURBINE SECTION OF HIGH BYPASS TURBOFAN

Номер: US20190017445A1
Автор: Adams Paul R., Lord Wesley K., Magge Shankar S., Schwarz Frederick M., Staubach Joseph B., Suciu Gabriel L.
Принадлежит:

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6. 1. A turbofan engine comprising:a fan including an array of fan blades rotatable about an engine axis;a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area;a fan duct including a fan duct annulus area outboard of the second compressor section inlet, wherein a ratio of the fan duct annulus area to the compressor inlet annulus area defines a bypass area ratio between 8.0 and 20.0;a turbine having a first turbine section driving the first compressor section, and a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and a second turbine airfoil count defined as the numerical count of all of the blades and vanes in the second turbine section;wherein a ratio of the second turbine airfoil count to the bypass area ratio is between 100 and 150;wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to ...

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Номер записи: 73
17-01-2019 дата публикации

TURBINE SECTION OF HIGH BYPASS TURBOFAN

Номер: US20190017446A1
Автор: Adams Paul R., Lord Wesley K., Magge Shankar S., Schwarz Frederick M., Staubach Joseph B., Suciu Gabriel L.
Принадлежит:

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, a shaft assembly having a first portion and a second portion, a turbine in fluid communication with the combustor, the turbine having a first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, an epicyclic transmission coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn faster than the fan, wherein the second turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

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Номер записи: 74
17-01-2019 дата публикации

VALVE SYSTEM FOR A FLUID CONDUIT SYSTEM IN AN AIRCRAFT ENGINE AND METHOD FOR THE OPERATION OF A VALVE SYSTEM FOR A FLUID CONDUIT SYSTEM IN AN AIRCRAFT ENGINE

Номер: US20190017524A1
Автор: FECHNER Stefan
Принадлежит:

A valve system for a fluid line system in an aircraft engine, which fluid line system has at least one fluid line wherein the at least one fluid line has at least one check valve, wherein the valve position of an actuator in the at least one check valve is changeable, in particular automatically changeable, in dependence on the pressure ratios in each case acting on the at least one check valve. The valve system has a monitoring means for recording the respective valve position, in particular the open position and/or the closed position of the actuator, in dependence on at least one measurement value, wherein a signal is output in dependence on the recorded valve position. The valve system furthermore has a means for setting a minimum required sealing-air stream. The invention also relates to a valve control method.

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Номер записи: 75
21-01-2021 дата публикации

Containment Case Active Clearance Control Structure

Номер: US20210017877A1
Автор: Jain Manoj Kumar, Jain Nitesh, Kray Nicholas Joseph
Принадлежит:

A propulsion system including a casing surrounding a fan rotor assembly is provided. An example casing includes an outer layer material, an inner layer material including first openings extended partially through the inner layer material along a radial direction of a first side of the inner layer material, and second openings extended partially through the inner layer material along the radial direction of a second side of the inner layer material opposite the first side, and a spring member coupled to the outer layer material and the inner layer material. 1. A casing to surround a fan rotor assembly , the casing comprising:an outer layer material; first openings extended partially through the inner layer material along a radial direction of a first side of the inner layer material; and', 'second openings extended partially through the inner layer material along the radial direction of a second side of the inner layer material opposite the first side; and, 'an inner layer material includinga spring member coupled to the outer layer material and the inner layer material.2. The casing of claim 1 , wherein the first openings and the second openings are configured to enable expansion and contraction of the inner layer material along the radial direction.3. The casing of claim 1 , wherein the outer layer material has a first coefficient of thermal expansion (CTE) and the spring member has a second CTE greater than the first CTE.4. The casing of claim 1 , wherein the spring member is disposed between the outer layer material and the inner layer material within a flow passage defined between an inner surface of the outer layer material and the first side of the inner layer material.5. The casing of claim 4 , wherein the spring member is extended at least partially along at least one of an axial direction or a circumferential direction within the flow passage.6. The casing of claim 1 , wherein the spring member defines a geometry based on a fin claim 1 , a ligament claim 1 ...

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Номер записи: 76
16-01-2020 дата публикации

GAS TURBINE ENGINE WITH ACTIVE VARIABLE TURBINE COOLING

Номер: US20200018240A1
Автор: Dvorak Steven M., Galle Nathan K.
Принадлежит:

A gas turbine engine includes a compressor section, a combustor section, and a turbine section operably coupled to the compressor section. A primary flow path is defined through the compressor section, the combustor section, and the turbine section. An engine case surrounds the compressor section, the combustor section, and the turbine section. The gas turbine engine also includes a means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section. 1. A gas turbine engine comprising:a compressor section;a combustor section;a turbine section operably coupled to the compressor section, wherein a primary flow path is defined through the compressor section, the combustor section, and the turbine section;an engine case surrounding the compressor section, the combustor section, and the turbine section; anda means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section.2. The gas turbine engine of claim 1 , wherein the means for providing the active variable cooling flow comprises a cooling air metering valve.3. The gas turbine engine of claim 2 , wherein the cooling air metering valve is electronically actuated based on either or both of a flight phase and an operating parameter of the gas turbine engine.4. The gas turbine engine of claim 1 , wherein the means for providing the active variable cooling flow comprises an airflow path within a static structure of the gas turbine engine between the bypass duct and the secondary flow cavity of the turbine section.5. The gas turbine engine of claim 4 , wherein the airflow path comprises an inner diffuser flow path configured to deliver a metered supply of cooling air from the compressor section through a diffuser section proximate to the combustor section.6. The gas turbine engine of claim 1 , wherein the secondary flow cavity of the turbine section ...

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Номер записи: 77
25-01-2018 дата публикации

ALTERNATING STARTER USE DURING MULTI-ENGINE MOTORING

Номер: US20180022465A1
Автор: Clauson Jesse W., Gelwan David, Schwarz Frederick M.
Принадлежит:

A system is provided for alternating starter use during multi-engine motoring in an aircraft. The system includes a first engine starting system of a first engine and a controller. The controller is operable to coordinate control of the first engine starting system to alternate use of power from a power source relative to use of the power by one or more other engine starting systems of the aircraft while maintaining a starting spool speed of the first engine below a resonance speed of the starting spool during motoring of the first engine. 1. A system for alternating starter use during multi-engine motoring in an aircraft , the system comprising:a first engine starting system of a first engine; anda controller operable to coordinate control of the first engine starting system to alternate use of power from a power source relative to use of the power by one or more other engine starting systems of the aircraft while maintaining a starting spool speed of the first engine below a resonance speed of the starting spool during motoring of the first engine.2. The system as in claim 1 , wherein control of the first engine starting system comprises intermittently accelerating and decelerating the starting spool of the first engine during motoring.3. The system as in claim 2 , wherein acceleration and deceleration of one or more other engines of the aircraft during motoring is performed in an alternating sequence with respect to the first engine.4. The system as in claim 1 , wherein timing of coordinated control of the first engine starting system and the one or more other engine starting systems is determined based on a plurality of performance parameters that are based on one or more of: an ambient condition claim 1 , performance limitations of a power source and each starter driven by the power source claim 1 , engine drag claim 1 , and parasitic factors.5. The system as in claim 4 , wherein the performance parameters are determined based on one or more of: an ambient air ...

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Номер записи: 78
28-01-2016 дата публикации

COMPRESSOR ASSEMBLY FOR GAS TURBINE

Номер: US20160024970A1
Автор: GOMEZ-LOR-LOPEZ Isaac-Eliseo, OLMES Sven, ZIERER Thomas
Принадлежит:

A compressor assembly, and more in general relates to a compressor for a gas turbine providing a solution that teaches to locate within a cavity formed by the outer casing of the compressor and the inner vane carrier a separator element, or membrane, such to divide the cavity into two sub-cavities. This advantageously results in a more flexible design with respect to the positioning of the flange blow-off extractor and to the cavity sizing, as the flange position is not necessarily the boundary for the flow anymore as it would be without the separator element. 1. A compressor assembly for a compressor of a gas turbine , the compressor assembly comprising:a compressor outer casing comprising at least one blow-off opening;a vane carrier defining a bleed duct;wherein the compressor assembly is arranged such that the outer casing and the vane carrier define a cavity for gathering a flow of fluid, said cavity being adapted to receive the fluid through said bleed duct and to feed the fluid externally through said blow-off opening; andwherein the compressor assembly includes a separator element located in said cavity such to divide said cavity in two sub-cavities.2. The compressor assembly according to claim 1 , wherein said separator element extends along a radial direction R of the compressor.3. The compressor assembly according to claim 1 , wherein said separator element is arranged between said outer casing and said vane carrier.4. The compressor assembly according to claim 1 , wherein at least an inner wall of the cavity is covered with a thermally insulating layer.5. The compressor assembly according claim 1 , wherein said thermally insulating layer comprises a coating material.6. The compressor assembly according to claim 1 , wherein said coating material is a ceramic-based coating.7. The compressor assembly according to claim 4 , wherein said thermally insulating layer comprises a metal sheet positioned on said inner wall of the cavity claim 4 , said inner wall ...

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Номер записи: 79
28-01-2016 дата публикации

HIGH TEMPERATURE DISK CONDITIONING SYSTEM

Номер: US20160025014A1
Автор: Forcier Matthew P., Schwarz Frederick M.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas-circulation system for conditioning a disk of an aircraft may comprise a first takeoff port configured to extract a combusted gas and a second takeoff port configured to extract an uncombusted gas. A first valve may comprise an inlet in fluid communication with the first and second takeoff ports and an outlet of the first valve in fluid communication with the disk. 1. A gas-circulation system for conditioning a disk in a gas turbine engine , the gas-circulation system comprising:a first takeoff port configured to extract a combusted gas;a second takeoff port configured to extract an uncombusted gas; anda first valve comprising an inlet in fluid communication with the first takeoff port and the second takeoff port and an outlet of the first valve in fluid communication with the disk.2. The gas-circulation system of claim 1 , further comprising a second valve fluidly coupled between the outlet of the first valve and the disk.3. The gas-circulation system of claim 2 , further comprising a first thermocoupling thermally coupled to a second valve outlet of the second valve.4. The gas-circulation system of claim 3 , further comprising a second thermocoupling thermally coupled to the second valve outlet.5. The gas-circulation system of claim 2 , wherein the second valve is configured to fail in a closed position.6. The gas-circulation system of claim 1 , further comprising:a combustor, wherein the first takeoff port is in fluid communication with the combustor;a turbine in fluid communication with and configured to be driven by the combusted gas from the combustor; anda compressor in fluid communication with and configured to supply the uncombusted gas to the combustor, wherein the second takeoff port is in fluid communication with the compressor.7. The gas-circulation system of claim 6 , wherein the first takeoff port extracts the combusted gas forward of the turbine.8. The gas-circulation system of claim 6 , wherein the first takeoff port extracts the combusted gas ...

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Номер записи: 80
26-01-2017 дата публикации

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

Номер: US20170022847A1
Автор: Eunkyeong Kim, Fumiyuki Suzuki, Kazunori Yamanaka, Miyuki Kawata, Norihiro Iyanaga, Takuya Yoshida, Tatsuro Yashiki, Yasuhiro Yoshida
Принадлежит: Mitsubishi Hitachi Power Systems Ltd

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.

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Номер записи: 81
26-01-2017 дата публикации

SIMPLIFIED ENGINE BLEED SUPPLY WITH LOW PRESSURE ENVIRONMENTAL CONTROL SYSTEM FOR AIRCRAFT

Номер: US20170022899A1
Автор: Schwarz Frederick M., Suciu Gabriel L.
Принадлежит:

A gas turbine engine assembly for an aircraft includes a fan section delivering air into a main compressor section. The main compressor section includes a first compressor section and a second compressor section operating at a higher pressure than the first compressor section. The main compressor section compresses air and delivers air into a combustion section where the air is mixed with fuel and ignited to generate products of combustion that are passed over a turbine section to drive the fan section and main compressor sections. An environmental control system includes a low pressure tap at a location on the first compressor section of the main compressor section. The low pressure tap communicates airflow to a first passage leading to a downstream outlet. A compressor is driven by an electric motor. A combined outlet intermixes airflow from the first passage and from the compressor driven by the electric motor and passes the airflow downstream to be delivered to an aircraft use. An environmental control system is also disclosed. 1. A gas turbine engine assembly for an aircraft comprising:a fan section delivering air into a main compressor section, said main compressor section including a first compressor section and a second compressor section operating at a higher pressure than the first compressor section, said main compressor section compressing air and delivering air into a combustion section where the air is mixed with fuel and ignited to generate products of combustion that are passed over a turbine section to drive said fan section and main compressor sections;an environmental control system including a low pressure tap at a location on the first compressor section of the main compressor section,wherein the low pressure tap communicates airflow to a first passage leading to a downstream outlet, and a compressor driven by an electric motor; anda combined outlet intermixing airflow from the first passage and from the compressor driven by the electric motor ...

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Номер записи: 82
26-01-2017 дата публикации

Low pressure compressor diffuser and cooling flow bleed for an industrial gas turbine engine

Номер: US20170022905A1
Автор: Brostmeyer Joseph D., Cejka Justin T., Jones Russell B., Orosa John A.
Принадлежит:

An industrial gas turbine engine with a high spool and a low spool in which low pressure compressed air is supplied to the high pressure compressor, and where a portion of the low pressure compressed air is bled off for use as cooling air for hot parts in the high pressure turbine of the engine. Annular bleed off channels are located in the LPC diffuser. The bleed channels bleed off around 15% of the core flow and pass the bleed off air into a cooling flow channel that then flows into the cooling circuits in the turbine hot parts. 1. An industrial gas turbine engine for electrical power production comprising:a high spool with a high pressure compressor and a high pressure turbine;a low spool with a low pressure compressor (LPC) and a low pressure turbine;a compressed air duct connecting a core flow of the LPC to an inlet of the high pressure compressor;a LPC diffuser air bleed channel to bleed off a portion of the core flow; and,a cooling flow channel connected to the LPC diffuser air bleed channel.2. The industrial gas turbine engine of claim 1 , and further comprising:the LPC diffuser air bleed channel bleeds off around 7.5% of the core flow of the LPC diffuser.3. The industrial gas turbine engine of claim 1 , and further comprising:a second LPC diffuser air bleed channel located downstream from the first LPC diffuser air bleed channel to bleed off a second portion of the core flow; and,the second LPC diffuser air bleed channel is connected to the cooling flow channel.4. The industrial gas turbine engine of claim 3 , and further comprising:the first and second LPC diffuser air bleed channels bleed off around 15% of the core flow of the LPC diffuser.5. The industrial gas turbine engine of claim 1 , and further comprising:the LPC diffuser air bleed channel is an annular shaped channel.6. The industrial gas turbine engine of claim 3 , and further comprising:the first and second LPC diffuser air bleed channels are both annular in shape; and,compressed air from the ...

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Номер записи: 83
25-01-2018 дата публикации

FINE DEBRIS MULTI-STAGE SEPARATION SYSTEM

Номер: US20180023473A1
Автор: MANTEIGA John Alan, MOOK Joshua Tyler, OTTA Shourya Prakash
Принадлежит:

The present disclosure generally relates to separating entrained solid particles from an input airflow in a gas turbine engine. A cyclonic separator receives the input airflow from a compressor and separates a first portion of the input airflow. The cyclonic separator remove solid particles from the first portion of the input airflow to provide a first cleaned airflow to a first cooling system. A clean air offtake downstream from the cyclonic separator separates a second cleaned airflow from a remaining portion of the input air stream and provides the second cleaned airflow to a second cooling system. The remaining portion of the input airflow is provided to a combustor. 1. A system for separating entrained solid particles from an input airflow in a gas turbine engine , comprising:a cyclonic separator that receives the input airflow from a compressor and separates a first portion of the input airflow, the cyclonic separator further removes solid particles from the first portion of the input airflow to provide a first cleaned airflow to a first cooling system; anda clean air offtake downstream from the cyclonic separator that separates a second cleaned airflow from a remaining portion of the input air stream and provides the second cleaned airflow to a second cooling system,wherein the remaining portion of the input airflow is provided to a combustor.2. The system of claim 1 , wherein the compressor includes a centripetal impeller.3. The system of claim 2 , wherein the cyclonic separator includes at least one opening arranged circumferentially about a compressor impeller shroud surrounding the centripetal impeller.4. The system of claim 3 , where each of the at least one opening is connected to a respective vortex chamber of the cyclonic separator.5. The system of claim 4 , wherein each vortex chamber includes a vortex finder that separates an outer vortex fed by the respective opening from an inner vortex of relatively clean air.6. The system of claim 5 , wherein ...

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Номер записи: 84
25-01-2018 дата публикации

GAS TURBINE ENGINE WITH HEAT PIPE FOR THERMAL ENERGY DISSIPATION

Номер: US20180023475A1
Автор: Xu JinQuan
Принадлежит:

A gas turbine engine includes a core nacelle and a core engine disposed in the core nacelle. The core engine includes at least a compressor section, a combustor section, and a turbine section, which define a core flowpath. A bypass duct is radially outwards of the core nacelle. A component is disposed in the core engine. A cooling cavity is disposed outside of the bypass duct. A heat pipe contains a working medium sealed therein. The heat pipe includes a first section configured to accept thermal energy and a second section configured to dissipate the thermal energy in the cooling cavity. A tap from at least one of the compressor section or the bypass duct is configured to provide bleed air to the cooling cavity and dissipate the thermal energy from the second section. 1. A gas turbine engine comprising:a core nacelle;a core engine disposed in the core nacelle, the core engine including at least a compressor section, a combustor section, and a turbine section, which define a core flow path;a bypass duct radially outwards of the core nacelle;a component disposed in the core engine;a cooling cavity disposed outside of the bypass duct;a heat pipe containing a working medium sealed therein, the heat pipe including a first section configured to accept thermal energy and a second section configured to dissipate the thermal energy in the cooling cavity; anda tap from at least one of the compressor section or the bypass duct, the tap configured to provide bleed air to the cooling cavity and dissipate the thermal energy from the second section.2. The gas turbine engine as recited in claim 1 , wherein the cooling cavity is disposed radially between the core flow path and the core nacelle.3. The gas turbine engine as recited in claim 1 , wherein the core nacelle defines an outer boundary of the cooling cavity.4. The gas turbine engine as recited in claim 1 , wherein the tap is from the compressor section.5. The gas turbine engine as recited in claim 1 , wherein the tap is from ...

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Номер записи: 85
25-01-2018 дата публикации

SHAFT ASSEMBLY OF A GAS TURBINE ENGINE AND METHOD OF CONTROLLING FLOW THEREIN

Номер: US20180023476A1
Автор: PARADIS Vincent, SYNNOTT Remy
Принадлежит:

A gas turbine engine comprises a shaft assembly including a hollow shaft of the gas turbine engine and a plug connected to the inlet end of the shaft. The hollow shaft has a shaft bore having a bore diameter. The hollow shaft has an inlet end for receiving a first portion of an incoming air flow. The plug has a plug bore therethrough, and an inlet end having an inlet diameter. The inlet diameter of the plug is smaller than the bore diameter. The plug includes a deflection surface adapted to deflect a second portion of the incoming air flow away from the shaft bore. A plug for connecting to an end of a hollow shaft of a gas turbine engine and s method of controlling a flow of fluid through a shaft having a bore therethrough of a gas turbine engine are also presented. 1. A plug for connecting to an end of a hollow shaft of a gas turbine engine , the plug comprising:a plug body having a first end and a second end, the first end being adapted to be connected to the hollow shaft;a bore extending through the body from the first end to the second end, the first end having a first inner diameter larger than a second inner diameter of the second end; anda flaring deflection surface disposed on an outer surface of the plug body between the first end and the second end.2. The plug as defined in claim 1 , wherein the first end of the plug is threaded.3. The plug as defined in claim 1 , wherein the bore includes a fairing portion extending from first end toward the second end.4. The plug as defined in claim 1 , further comprising a grasping portion disposed between the first end and the deflection surface.5. The plug as defined in claim 4 , further comprising the grasping portion is hexagonal. The present application is a divisional application of U.S. patent application Ser. No. 14/230,789 filed on Mar. 31, 2014, incorporated herein by reference.The application relates generally to gas turbine engines and, more particularly, to cooling of shafts of gas turbine engines.Gas ...

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Номер записи: 86
25-01-2018 дата публикации

PASSIVE HEAT EXCHANGER VALVE

Номер: US20180023906A1
Автор: Jonnalagadda Dattu GV, Kenworthy Michael Thomas, Tajiri Gordon, Yang Yanzhe
Принадлежит:

A valve and a heat exchanger apparatus for a gas turbine engine include a valve body having a valve seat and an actuation component including a plate formed from a set of metal layers and responsive to a change in at least one of a thermal condition and a pressure exerted thereon such that the plate moves and the valve moves between an opened and a closed position where a portion of the plate engages with the valve seat. 1. A valve , comprising:a valve body having a valve seat; andan actuation component including a plate formed from a set of metal layers responsive to a change in at least one of a thermal condition or a pressure exerted thereon such that the valve moves between an opened and a closed position where the plate engages with the valve seat.2. The valve of wherein the plate is mounted to the valve body at a first end and a second end of the plate defines a free end.3. The valve of wherein the plate comprises an extension at its free end that is configured to engage with the valve seat of the valve body.4. The valve of wherein the plate is configured to move the extension to the closed position in response to a predetermined temperature exerted thereon.5. The valve of wherein the plate includes a monolithic composite metal sheet.6. The valve of wherein the plate includes a bi-metal strip comprising a layer of aluminum alloy and a layer of aluminum silicon carbide.7. The valve of wherein the coefficient of thermal expansion for the layer of aluminum silicon carbide ranges from 20 ppm/C to 7 ppm/C.8. The valve of wherein the valve is incorporated in a heat exchanger.9. The valve of wherein the heat exchanger is an air-cooled oil cooler for a turbine engine.10. The valve of wherein the plate comprises two fixed ends and a central extension configured to engage the valve seat of the valve body.11. The valve of wherein the plate further comprises a secondary temperature controlled portion that is configured to allow bi-modal flows.12. The valve of wherein the ...

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Номер записи: 87
24-04-2014 дата публикации

Part load performance improvement using deformable bore plugs

Номер: US20140109580A1
Автор: Bhaskar Pemmi, Harish BOOMANAKATTE, Indrajit Mazumder, Krishna Kishore Gumpina, Rajarshi Saha, Sanjeev Kumar Jha, Santhosh Donkada, Sheo Narain Giri
Принадлежит: General Electric Co

A cooling arrangement for a gas turbine engine. The cooling arrangement comprises a discharge channel for air flow from a compressor, a first cooling channel and at least one aperture providing communication between the flow of air through the discharge channel and the first cooling channel. A restrictor device in the aperture regulates the flow of air between the discharge channel and the first cooling channel. The restrictor device deforms to vary air flowing through the aperture in response to a physical condition of the engine. This physical condition of the engine may be that of the temperature of air flowing through the discharge channel, the restrictor device responding to regulate the flow of air based on that temperature. The restrictor device may be a two-way shape memory alloy.

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Номер записи: 88
24-01-2019 дата публикации

GAS TURBINE ENGINE WITH ROTOR TIP CLEARANCE CONTROL SYSTEM

Номер: US20190024527A1
Автор: Eifert Andrew J., Skertic Richard J.
Принадлежит:

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 ...

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Номер записи: 89
24-01-2019 дата публикации

REVERSE CORE GEAR TURBOFAN

Номер: US20190024584A1
Автор: Kupratis Daniel Bernard, Moon Francis R.
Принадлежит:

A gas turbine engine comprises a fan at an axially outer location, the fan rotating about an axis of rotation, delivering air into an outer bypass duct, a radially middle duct, and a radially inner core duct. Air from the inner core duct is directed into a compressor, and then flows axially in a direction back toward the fan through a combustor section, and across a core turbine section, and is then directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method of operating a gas turbine engine is also disclosed. 1. A gas turbine engine comprising:a fan at an axially outer location, said fan rotating about an axis of rotation;said fan delivering air into an outer bypass duct, a radially middle duct, and a radially inner core duct;air from said inner core duct being directed into a compressor, and then flowing axially in a direction back toward said fan through a combustor section, and across a core turbine section, and then being directed into said middle duct;a gear reduction for driving said fan from a fan drive turbine section; andwherein said cold turbine section is in said radially inner core duct and is provided with a flow diverter that allows bypass of air around a rotor associated with said cold turbine section.2. The gas turbine engine as set forth in claim 1 , wherein a shaft downstream of said gear reduction relative to said fan drive turbine section also drives said booster fan.3. The gas turbine engine as set forth in claim 2 , wherein a shaft downstream of said gear reduction is also connected to rotate with said cold turbine.4. The gas turbine engine a set forth in claim 3 , wherein a fan booster delivering air into said radially middle duct and said radially inner core claim 3 , and a cold turbine in said radially inner core rotating with a clutched shaft separate from a fan shaft driving said fan claim 3 , and a clutch selectively connecting said clutched shaft to said fan shaft such that said fan shaft ...

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Номер записи: 90
24-01-2019 дата публикации

FAN INTEGRATED INERTIAL PARTICLE SEPARATOR

Номер: US20190024587A1
Автор: Bryan William B., III Crawford F., Lerg Bryan H., Oechsle Victor, Smith
Принадлежит:

A gas turbine engine includes a fan, an engine core, and an airflow duct assembly. The fan is mounted for rotation about a central axis of the gas turbine engine assembly to produce thrust for the gas turbine engine. The engine core is coupled to the fan and configured to drive the fan about the central axis. The airflow duct assembly defines a core passageway configured to conduct a first portion of air pushed by the fan into the engine core and a by-pass passageway configured to conduct a second portion air pushed by the fan around the engine core. 1. A gas turbine engine comprisinga fan mounted for rotation about a central axis of the gas turbine engine,an engine core coupled to the fan and configured to drive the fan about the central axis to cause the fan to push a mixture of air and particles suspended in the air to provide thrust for the gas turbine engine, andan airflow duct assembly configured to conduct the mixture of air and particles through the gas turbine engine, the airflow duct assembly defining a core passageway configured to conduct a first portion of the mixture of air and particles pushed by the fan into the engine core and a by-pass passageway configured to conduct a second portion of the mixture of air and particles pushed by the fan around the engine core, andwherein the airflow duct assembly includes a particle-separator splitter positioned in the core passageway and configured to separate the first portion of the mixture of air and particles into a clean flow substantially free of particles and a dirty flow containing the particles and the particle-separator splitter is arranged to direct the clean flow into the engine core and the dirty flow away from the engine core.2. The gas turbine engine of claim 1 , wherein the airflow duct assembly further includes an inner wall arranged circumferentially around the central axis claim 1 , an outer wall arranged circumferentially around the inner wall and the fan claim 1 , and a by-pass flow splitter ...

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Номер записи: 91
24-01-2019 дата публикации

COMPACT AERO-THERMO MODEL STABILIZATION WITH COMPRESSIBLE FLOW FUNCTION TRANSFORM

Номер: US20190024590A1
Автор: Karpman Boris, Shishkin Serge L.
Принадлежит:

Systems and methods for controlling a fluid based engineering system are disclosed. The systems and methods may include a model processor for generating a model output, the model processor including a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode, where the open loop model generates current state derivatives, solver state errors, and synthesized parameters as a function of the dynamic states and a model input vector, where a constraint on the current state derivatives and solver state errors is based a series of cycle synthesis modules. The series of cycle synthesis modules may include a flow module for mapping a flow curve relating a compressible flow function to a pressure ratio and for defining a solution point located on the flow curve and a base point located off the flow curve. 1. A control system , comprising:an actuator configured to position a control device, the control device defining a flow path through an aperture, the aperture defining a pressure drop along the flow path, and comprising a control surface, wherein the actuator positions the control surface in order to regulate fluid flow through the aperture;a control law configured to direct the actuator as a function of a model output; and an input object for processing a model input vector and setting a model operating mode;', 'a set state module for setting dynamic states of the model processor, the dynamic states input to an open loop model based on the model operating mode;', 'wherein the open loop model generates current state derivatives, solver state errors, and synthesized parameters as a function of the dynamic states and the model input vector, wherein a constraint on the current state derivatives and solver state errors is based on a series of cycle synthesis modules, each member of the series of cycle synthesis modules modeling a component of a cycle of the control device and comprising a ...

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Номер записи: 92
24-01-2019 дата публикации

GEARED TURBOFAN GAS TURBINE ENGINE ARCHITECTURE

Номер: US20190024610A1
Автор: Houston David P., Kupratis Daniel Bernard, Schwarz Frederick M.
Принадлежит:

A gas turbine engine according to an example of the present disclosure includes a turbine section having a very high speed fan drive turbine such that a quantity defined by the exit area of the fan drive turbine multiplied by the square of the fan drive turbine rotational speed compared to the same parameters for a second turbine is at a ratio between 0.5 and 1.5. The engine includes a power density greater than 1.5 lbf/in. 1. A gas turbine engine comprising:a fan section including a fan rotor and fan, the fan including a plurality of fan blades, with a low fan pressure ratio in operation of less than 1.45, the low fan pressure ratio measured across the fan blades alone;a compressor section;wherein the fan delivers a portion of air into a bypass duct, and a bypass ratio being defined as the portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section, and the bypass ratio is greater than 10;a gear train having a gear ratio greater than 2.3;a combustor in fluid communication with the compressor section; [{'sup': '3', 'the engine includes a power density greater than 1.5 lbf/in, and power density is defined as Sea Level Takeoff Thrust in lbf produced divided by the volume of the turbine section in cubic inches;'}, 'the fan drive turbine includes at least one rotor having a bore radius (R), a live rim radius (r) and a bore width (W), and a ratio of r/R is between 2.00 and 2.30, and a ratio of r/W is between 4.65 and 5.55;', 'the fan drive turbine includes an inlet, an outlet, and a fan drive turbine pressure ratio greater than 5 in operation, wherein the fan drive turbine pressure ratio is a ratio of a pressure measured prior to the inlet as related to a pressure at the outlet prior to any exhaust nozzle;', 'the fan drive turbine has a first exit area and rotates at a first speed, the second turbine has a second exit area and rotates at a second speed, which is faster than the first speed, wherein a first performance ...

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Номер записи: 93
23-01-2020 дата публикации

ANTI-ICE SYSTEM EXHAUST AIR DISRUPTOR

Номер: US20200025072A1
Автор: Mackin Steve G., Miller Kimberly A.
Принадлежит:

An airflow disruptor for a gas turbine engine bleed air exhaust port employs a disruptor plate rotatably mounted upstream from an exhaust port of an exhaust duct. An actuator is coupled to the disruptor plate and adapted to rotate the disruptor plate into an external airflow responsive to temperature of exhaust flow in the exhaust port whereby the external airflow is turbulated upstream of the exhaust port. 1. An airflow disruptor for a gas turbine engine bleed air exhaust port , the airflow disruptor comprising:a disruptor plate rotatably mounted upstream from an exhaust port of an exhaust duct; and,an actuator coupled to the disruptor plate, the actuator adapted to rotate the disruptor plate into an external airflow responsive to temperature of exhaust flow in the exhaust port whereby the external airflow is turbulated upstream of the exhaust port.2. The airflow disruptor as defined in further comprising a temperature sensor and wherein the actuator is operable responsive to a temperature signal from the temperature sensor.3. The airflow disruptor as defined in further comprising a control system receiving the temperature signal and issuing an actuation signal to the actuator responsive to the temperature signal exceeding a threshold value.4. The airflow disruptor as defined in wherein the actuation signal is proportional to temperature over the threshold value as indicated by the temperature signal.5. The airflow disruptor as defined in wherein the disruptor plate is rotatable through a range of positions from fully closed to fully open responsive to the temperature signal.6. The airflow disruptor as defined in wherein the disruptor plate incorporates perforations.7. The airflow disruptor as defined in wherein the disruptor plate incorporates a trailing edge treatment.8. The airflow disruptor as defined in wherein the actuator is a thermal actuator claim 1 , said thermal actuator positioned within the exhaust duct.9. The airflow disruptor as defined in wherein ...

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Номер записи: 94
23-01-2020 дата публикации

GAS TURBINE ENGINE

Номер: US20200025078A1
Автор: CULWICK Steven P., PERERA Shane, SCOTHERN David P., SPALTON Edward J.
Принадлежит: ROLLS-ROYCE PLC

A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core. 1. A gas turbine engine comprising:an engine core comprising a compressor;a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor;at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; anda flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core;wherein the flow controller is configured to provide a first mass flow rate of bleed air when the bleed air is provided in response to a requirement to de-ice the component and a second mass flow rate of bleed air, different from the first mass flow rate, when the bleed air is provided in response to a requirement to release bleed air from the compressor to optimise operation of the core, and to provide the higher of the first and second mass flow rates of bleed air when the bleed air is required both for de-icing the component and to optimise operation of the core.2. The gas turbine engine of claim 1 , further comprising an icing detector configured to detect a level of icing on the at least one component; wherein the flow controller is configured to provide a mass flow rate of bleed air to the de- ...

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Номер записи: 95
23-01-2020 дата публикации

DIVERSION OF FAN AIR TO PROVIDE COOLING AIR FOR GAS TURBINE ENGINE

Номер: US20200025080A1
Автор: Bifulco Anthony R., Kupratis Daniel Bernard
Принадлежит:

A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages. 1. A gas turbine engine section comprising:a plurality of spaced rotor stages, with a static guide vane intermediate said spaced rotor stages, said static guide vane providing swirl into air passing toward a downstream one of said spaced rotor stages, and an outer housing surrounding said spaced rotor stages, a diverter diverting a portion of air radially outwardly through said outer housing, and across at least one heat exchanger, with the diverted air passing back into a duct radially inwardly through said outer housing, and being exhausted toward said downstream one of said spaced rotor stages.2. The gas turbine engine as set forth in claim 1 , wherein said exhausted air passing through an injector claim 1 , and said injector imparting swirl into the air exhausting toward the downstream one of the two spaced turbine rotors.3. The gas turbine engine as set forth in claim 2 , wherein a swirl angle imparted by said injector is greater than a swirl angle imparted by said static guide vane.4. The gas turbine engine as set forth in claim 3 , wherein a swirl angle imparted by said static guide vane is greater than 40 degrees.5. The gas turbine engine as set forth in claim 4 , wherein said swirl angle is greater than 55 degrees.6. The gas turbine engine as set forth in claim 1 , wherein said at least one heat exchanger cooling an electronic component.7. The gas turbine engine as set forth in claim 1 , ...

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Номер записи: 96
23-01-2020 дата публикации

SELF-REGULATING BACK-SIDE PRESSURIZATION SYSTEM FOR THERMAL INSULATION BLANKETS

Номер: US20200025091A1
Автор: De Pau, JR. Robert, Miller Kimberly A.
Принадлежит: The Boeing Company

High-pressure fan duct bleed air is used to pressurize a cavity between the fan duct inner wall and the inner wall thermal insulation blankets. The cavity is pressurized to prevent hot air from the nacelle core compartment from flowing under the insulation blankets and degrading the fan duct inner wall. Pressure regulating valves (PRV) allow better control of the cavity pressure during different phases of the flight profile and under different levels of insulation blanket seal degradation by passively controlling exit area from the cavity based on an established pressure limit. Moreover, the pressurization system can be implemented as a passive cooling system by increasing the mass flow rate into the cavity and then the core compartment to a level suitable for core compartment cooling. The cooling air can be vented at the forward end of the insulation blanket assembly to provide core compartment ventilation flow, or vented through dedicated ports in the insulation blanket for targeted core compartment component cooling. 1400. An aircraft engine () , comprising:{'b': '402', 'an engine core ();'}{'b': 404', '202, 'a fan duct () including an inner wall ();'}{'b': 1', '202, 'a first orifice (A) through the inner wall ();'}{'b': 206', '202', '202', '402, 'an insulation blanket () coupled to the inner wall () so as to shield the inner wall () from heat generated in the engine core ();'}{'b': 2', '206, 'a second orifice (A) through the insulation blanket ();'}{'b': 206', '202', '206', '206', '406', '206', '1', '404, 'a cavity () bounded by the inner wall () and the insulation blanket (), the cavity () receiving air () inputted into the cavity () through the first orifice (A) from the fan duct ();'}{'b': 210', '404', '402', '210', '408', '402', '410', '206, 'a core compartment () within the fan duct () and housing the engine core (), the core compartment () having a first boundary () with the engine core () and second boundary () with the insulation blanket (); and'}{'b': ...

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Номер записи: 97
23-01-2020 дата публикации

MULTI-BYPASS STREAM GAS TURBINE ENGINE WITH ENLARGED BYPASS FLOW AREA

Номер: US20200025104A1
Автор: Snape Nathan, Thomas Michael R.
Принадлежит:

A gas turbine engine comprises a first bypass flow path housing configured within the engine, radially exterior to an engine core housing, and a second bypass flow path housing configured within the engine, radially exterior to the first bypass flow path housing. An axially downstream portion of the first bypass flow path housing includes a stepwise increase in area compared with an axially adjacent upstream portion of the first bypass flow path housing, thereby defining a component placement cavity in the axially downstream portion. 1. A gas turbine engine comprising:a first bypass flow path housing that is radially exterior to an engine core housing with respect to a central longitudinal axis of the gas turbine engine;a second bypass flow path housing that is radially exterior to said first bypass flow path housing with respect to the central longitudinal axis of the engine;a core engine flow path defined by the engine core housing, the core engine flow path radially inward of the first bypass flow path housing with respect to the central longitudinal axis of the gas turbine engine;a first fan rotor that delivers air into said first bypass flow path housing, said second bypass flow path housing, and said core engine housing;wherein an axially downstream portion of said first bypass flow path housing includes a stepwise increase in area radially outwardly with respect to the central longitudinal axis compared with an axially adjacent upstream portion of said first bypass flow path housing, thereby defining a component placement cavity in said axially downstream portion that is circumferentially covered by a portion of the first bypass flow path housing;wherein the first bypass flow path housing includes at least one duct which extends radially outwardly with respect to the central longitudinal axis into a flow path within said second bypass flow path housing to define the stepwise increase in area, wherein the stepwise increase in area of each duct of said at least ...

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Номер записи: 98
23-01-2020 дата публикации

Intercooled cooling air with auxiliary compressor control

Номер: US20200025105A1
Автор: Brian D. Merry, Gabriel L. Suciu, Michael G. Mccaffrey
Принадлежит: United Technologies Corp

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.

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Номер записи: 99
23-01-2020 дата публикации

GEARED TURBOFAN WITH HIGH GEAR RATIO AND HIGH TEMPERATURE CAPABILITY

Номер: US20200025106A1
Автор: Schwarz Frederick M., Suciu Gabriel L.
Принадлежит:

A gas turbine engine comprises a fan rotor having blades with an outer diameter. The outer diameter is greater than or equal to 77 inches (196 centimeters) and less than or equal to 135 inches (343 centimeter). The fan rotor has less than or equal to 26 fan blades, and is driven by a fan drive turbine through a gear reduction. The gear reduction has a gear ratio of greater than 2.6:1. The fan rotor delivers air into a bypass duct as bypass air, and into a duct leading to a compressor rotor as core air. A ratio of bypass air to the core air is greater than or equal to 12:1. An upstream turbine rotor is upstream of the fan drive turbine and drives a compressor rotor. The upstream turbine rotor has at least two stages, and the fan drive turbine rotor has at least three stages. The turbine blades in at least one stage of the fan drive turbine rotor are provided with a performance enhancing feature, which is at least one of the blades being manufactured by a directionally solidified blade material. The blades are provided as single crystal blades, and have a radially outer platform provided with scalloping to reduce the weight of the blades. The blades are provided with cooling air. 1. A gas turbine engine comprising:a fan rotor having blades with an outer diameter, said outer diameter being greater than or equal to 77 inches (196 centimeters) and less than or equal to 135 inches (343 centimeter), and said fan rotor having less than or equal to 26 fan blades;said fan rotor driven by a fan drive turbine through a gear reduction, said gear reduction having a gear ratio of greater than 2.6:1, and said fan rotor delivering air into a bypass duct as bypass air, and into a duct leading to a compressor rotor as core air, and a ratio of bypass air to said core air being greater than or equal to 12:1;an upstream turbine rotor upstream of said fan drive turbine and driving a compressor rotor, and said upstream turbine rotor having at least two stages, and said fan drive turbine ...

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