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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 424. Отображено 159.
08-08-2017 дата публикации

Load balanced journal bearing pin for planetary gear

Номер: US0009726083B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A disclosed fan drive gear system includes a sun gear rotatable about an axis of rotation, a plurality of intermediate gears rotatable about an intermediate gear rotation axis in meshing engagement with the sun gear and a ring gear circumscribing the intermediate gears. A bearing assembly supports at least one of the plurality of intermediate gears and includes a first beam extending in a first direction and a second beam extending from an end of the first beam in a second direction. The bearing surface supported on the second beam such that first and second beams are configured to maintain the bearing surface substantially parallel to the intermediate gear rotation axis during operation.

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Номер записи: 1
18-08-2016 дата публикации

TURBINE ENGINE WITH A TURBO-COMPRESSOR

Номер: US20160237894A1
Автор: Daniel B. Kupratis, William G. Sheridan, KUPRATIS DANIEL B, SHERIDAN WILLIAM G, Kupratis Daniel B., Sheridan William G.
Принадлежит:

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

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Номер записи: 2
13-06-2017 дата публикации

Planetary gear system arrangement with auxiliary oil system

Номер: US0009677420B2
Автор: Michael E. McCune, William G. Sheridan, Lawrence E. Portlock, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, PORTLOCK LAWRENCE E, McCune Michael E., Sheridan William G., Portlock Lawrence E.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A gas turbine engine includes a fan, a speed reduction device driving the fan and a lubrication system for lubricating components across a rotation gap. The lubrication system includes a lubricant input. A stationary first bearing receives lubricant from the lubricant input and has a first race in which lubricant flows and a second race. A second bearing for rotation is within the first bearing including a first opening in registration with the first race such that lubricant may flow from the first race through the first opening into a first conduit. There is a rotating carrier for supporting at least one planetary gear. The second bearing extends from the rotating carrier about an axis. A first spray bar is disposed on the carrier. The second bearing has a second opening in registration with the second race and a second conduit for passing lubricant to the spray bar.

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Номер записи: 3
04-08-2016 дата публикации

FAN DRIVE GEAR SYSTEM

Номер: US20160222888A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section and a star gear system for driving the fan section. A first fan section support bearing is mounted forward of the star gear system and a second fan section bearing is mounted aft of the star gear system.

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Номер записи: 4
26-12-2017 дата публикации

Scavenge filter system for a gas turbine engine

Номер: US0009849411B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A scavenge filter system according to an exemplary aspect of the present disclosure includes, among other things, a first scavenge pump stage positioned in a first flow path downstream of a first bearing compartment of a spool and a second scavenge pump stage positioned in a second flow path downstream of a second bearing compartment. The second bearing compartment houses a geared architecture mechanically coupled to the spool. A first scavenge filter fluidly couples the first scavenge pump stage to at least one oil reservoir. A second scavenge filter fluidly couples the second scavenge pump stage to the at least one oil reservoir. The first and second scavenge filters are separate and distinct. A method of filtering debris is also disclosed.

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Номер записи: 5
29-12-2016 дата публикации

JOURNAL BEARING FOR ROTATING GEAR CARRIER

Номер: US20160377166A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gear system for a geared turbofan engine includes a sun gear, a planet gear supported in a carrier and engaged to the sun gear, a forward journal bearing and an aft journal bearing both supporting the planet gear. The carrier includes a forward wall supporting the forward journal bearing and an aft wall supporting the aft journal bearing. A ring gear is engaged to the planet gear. A geared turbofan engine is also disclosed.

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Номер записи: 6
23-02-2017 дата публикации

METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE

Номер: US20170051677A1
Автор: William G. Sheridan, Karl L. Hasel, SHERIDAN WILLIAM G, HASEL KARL L, Sheridan William G., Hasel Karl L.
Принадлежит:

A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram ° R)/(518.7° R)], where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio. A low and high pressure turbine in communication with a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages. 1. A gas turbine engine comprising:an engine centerline longitudinal axis;{'sup': '0.5', 'a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis, wherein the fan has a low corrected fan tip speed less than about 1400 ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram ° R)/(518.7° R)], where T represents the ambient temperature in degrees Rankine;'}a bypass ratio greater than about 11.0;a fan pressure ratio less than 1.48, wherein the fan pressure ratio is measured across a fan blade alone;a speed reduction device comprising a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1;a low pressure turbine in communication with a first shaft; anda high pressure turbine in communication with a second shaft;wherein the first shaft and second shaft are concentric and mounted via at least one of the plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is ...

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Номер записи: 7
17-08-2017 дата публикации

AUXILIARY DRIVE BOWED ROTOR PREVENTION SYSTEM FOR A GAS TURBINE ENGINE THROUGH AN ENGINE ACCESSORY

Номер: US20170234167A1
Автор: David Allen Stachowiak, William G. Sheridan, STACHOWIAK DAVID ALLEN, SHERIDAN WILLIAM G, Stachowiak David Allen, Sheridan William G.
Принадлежит:

A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a gear train and a bowed rotor prevention motor. The gear train is coupled through an engine accessory to an engine accessory gearbox that is further coupled to a starting spool of the engine. The bowed rotor prevention motor is operable to drive rotation of the starting spool of the gas turbine engine through the gear train upon engine shutdown.

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Номер записи: 8
17-08-2017 дата публикации

AUXILIARY DRIVE BOWED ROTOR PREVENTION SYSTEM FOR A GAS TURBINE ENGINE

Номер: US20170234232A1
Автор: William G. Sheridan, David Allen Stachowiak, SHERIDAN WILLIAM G, STACHOWIAK DAVID ALLEN, Sheridan William G., Stachowiak David Allen
Принадлежит:

A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a gear train and a bowed rotor prevention motor operable to drive rotation of a starting spool of the gas turbine engine through the gear train at a substantially constant speed upon engine shutdown until an energy storage source is depleted.

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Номер записи: 9
06-04-2017 дата публикации

GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20170096943A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an axis;a fan pressure ratio across a fan blade alone of less than 1.45;a bypass duct;a compressor section;a bypass ratio greater than ten (10), the 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;a combustor in fluid communication with the compressor section;a fan drive turbine in communication with the combustor, the fan drive turbine comprising a pressure ratio greater than about five (5), wherein the fan drive turbine further includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the fan drive turbine is a ratio of the inlet pressure to the outlet pressure;a gear system including a plurality of gears and configured to provide a speed reduction between the fan drive turbine and the fan and transfer power input from the fan drive turbine to the fan at an efficiency greater than 98%, wherein the gear system further comprises a gear reduction ratio of greater than 2.3;a gear support system including a spring rate that provides a defined amount of deflection and misalignment among at least some of the plurality of gears of the gear system;a lubrication system configured to provide lubricant to the gear system and remove thermal energy from the gear system; andwherein the plurality of fan blades is less ...

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Номер записи: 10
29-12-2016 дата публикации

LUBRICANT DELIVERY SYSTEM FOR PLANETARY FAN DRIVE GEAR SYSTEM

Номер: US20160377165A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gear system for a turbofan engine assembly includes a sun gear rotatable about an engine centerline, a non-rotatable ring gear, a rotating carrier that drives a fan, and a plurality of planet gears intermeshed between the sun gear and the ring gear. Each of the plurality of planet gears supported on rolling element bearings fit into the carrier. Each of the plurality of planet gears includes an inner cavity and a lubricant passage directed at the rolling element bearings. The carrier includes an outer scoop that receives lubricant from an outer fixed lubricant jet and feeds lubricant into the inner cavity and through the lubricant passage to spray lubricant on to the rolling element bearings. A geared turbofan engine assembly is also disclosed.

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Номер записи: 11
27-04-2017 дата публикации

GAS TURBINE ENGINE WITH GEARBOX HEALTH FEATURES

Номер: US20170114661A1
Автор: Frederick M. Schwarz, William G. Sheridan, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, Schwarz Frederick M., Sheridan William G.
Принадлежит:

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan and a braking system. The braking system is configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan. Further, when the rotation of the fan sufficiently slows, the braking system is further configured to apply a second level of braking more restrictive than the first level of braking. 1. A gas turbine engine , comprising:a fan; anda braking system configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan and, when the rotation of the fan sufficiently slows, apply a second level of braking more restrictive than the first level of braking.2. The gas turbine engine as recited in claim 1 , further comprising a gear reduction between the fan and a shaft of the engine.3. The gas turbine engine as recited in claim 1 , wherein the braking system includes a brake claim 1 , and wherein the fan includes a disc claim 1 , the brake configured to selectively engage the disc to apply the first level of braking to slow rotation of the fan.4. The gas turbine engine as recited in claim 3 , wherein the braking system further includes a pawl claim 3 , and wherein the disc includes a slot claim 3 , the pawl configured to selectively engage the slot to apply the second level of braking and to substantially lock the fan against rotation.5. The gas turbine engine as recited in claim 4 , wherein the brake engages the disc following an engine shut-off event and slows the rotation of the fan claim 4 , and claim 4 , when the rotation of the fan sufficiently slows claim 4 , the pawl engages the slot to lock the fan against further rotation.6. The gas turbine engine as recited in claim 4 , wherein the pawl is biased into engagement with the slot claim 4 , and wherein the pawl is selectively retracted from the slot by an actuator.7. The gas ...

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Номер записи: 12
09-08-2016 дата публикации

Auxiliary oil system for negative gravity event

Номер: US0009410448B2
Автор: William G. Sheridan, Paul H. Dolman, SHERIDAN WILLIAM G, DOLMAN PAUL H, Sheridan William G., Dolman Paul H.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, SHERIDAN WILLIAM G, DOLMAN PAUL H, UNITED TECHNOLOGIES CORP, Sheridan William G., Dolman Paul H.

A lubrication system includes a bearing compartment. A main reservoir is fluidly connected to the bearing compartment by a main supply passage. A main pump is arranged in the main supply passage configured to provide fluid from the main reservoir to the bearing compartment during a positive gravity condition. A secondary supply passage fluidly connects the main reservoir to at least one segment of the main supply passage, thereby providing fluid from the main reservoir to the bearing compartment during a negative gravity condition. A method of supplying a bearing compartment with fluid includes pumping a fluid from a main reservoir to a bearing compartment through a main supply passage during a positive gravity condition, and providing fluid from the main reservoir to the bearing compartment through a secondary supply passage, fluidly connected to at least one segment of the main supply passage, in response to a negative gravity condition.

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Номер записи: 13
06-07-2017 дата публикации

TRANSMISSION FOR COAXIAL MULTI-ROTOR SYSTEM

Номер: US20170190415A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A coaxial, dual rotor system includes a first rotor assembly located at a rotor axis and rotatable thereabout and a second rotor assembly located at the rotor axis radially inboard of the first rotor assembly and rotatable thereabout. A rotationally fixed static mast is located radially between the first rotor assembly and the second rotor assembly. A transmission includes a rotor input shaft including a first sun gear and a second sun gear. A star gear arrangement is operably connected to the first sun gear and to the static mast to drive rotation of the first rotor assembly about the rotor axis and a planetary gear arrangement is operably connected to the second sun gear and to the static mast to drive rotation of the second rotor assembly about the rotor axis.

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Номер записи: 14
30-01-2018 дата публикации

Oil loss protection for a fan drive gear system

Номер: US0009879608B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A fan drive gear system includes at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating. At least a first axial end of the intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system. The fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event.

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Номер записи: 15
12-09-2017 дата публикации

Oil baffle for gas turbine fan drive gear system

Номер: US0009759309B2
Автор: William G. Sheridan, Michael E. McCune, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, Sheridan William G., McCune Michael E.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

An epicyclic gear train component includes spaced apart walls with circumferentially spaced mounts that interconnect the walls. The mounts provide circumferentially spaced apart apertures between the mounts at an outer circumference of the walls. Baffles are arranged between the walls near the mounts. Gear pockets are provided between the baffles and the baffles include a lubrication passage that terminates at least one of the gear pockets. One of the walls includes a hole which has a tube that extends through the hole and is received in an opening in the baffle. The tube is in communication with the lubrication passage.

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Номер записи: 16
20-02-2018 дата публикации

Zero or low leakage oil transfer bearing

Номер: US0009896969B2
Автор: William G. Sheridan, Marc J. Muldoon, Michael E. McCune, SHERIDAN WILLIAM G, MULDOON MARC J, MCCUNE MICHAEL E, Sheridan William G., Muldoon Marc J., McCune Michael E.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A transfer bearing assembly is configured to allow a flow of a fluid from an inlet tube to a channel defined by a rotating shaft. The transfer bearing assembly includes a body having an axially forward side and an axially aft side and a first wing having an axially aft end coupled to the axially forward side and an axially forward end. The transfer bearing assembly also includes a first side plate having a radially outward end coupled to the axially forward end of the first wing and a radially inward end.

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Номер записи: 17
29-12-2016 дата публикации

ROLLER BEARINGS FOR HIGH RATIO GEARED TURBOFAN ENGINE

Номер: US20160376911A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gear system for a geared turbofan engine is disclosed. The gear system includes a sun gear driven by a low spool shaft. The sun gear defines a sun gear diameter. A rotating carrier drives a fan. The carrier defines an outer carrier diameter and an inner carrier diameter. A non-rotating ring gear is also included. The ring gear defines a ring gear diameter and the ring gear diameter is smaller than the outer carrier diameter. A set of planet gears are mounted on corresponding rolling element bearing assemblies. Each roller element bearing assembly is supported within the carrier within a space defined between the carrier outer diameter and the carrier inner diameter. Each of the sun gear, ring gear and planet gears are substantially centered along a gearbox centerline transverse to an engine longitudinal axis and the gear system provides a speed reduction ratio between an input to the sun gear and an output from the carrier between 3:1 and 5:1. A method of creating a gear system for a ...

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Номер записи: 18
06-04-2017 дата публикации

GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20170096937A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an axis;a fan pressure ratio across a fan blade alone of less than 1.45;a bypass duct;a compressor section;a bypass ratio greater than ten (10), the 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;a combustor in fluid communication with the compressor section;a fan drive turbine in communication with the combustor, the fan drive turbine comprising a pressure ratio greater than about five (5), wherein the fan drive turbine further includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the fan drive turbine is a ratio of the inlet pressure to the outlet pressure;a gear system configured to provide a speed reduction between the fan drive turbine and the fan and transfer power input from the fan drive turbine to the fan at an efficiency greater than 98%, wherein the gear system further comprises a gear reduction ratio of greater than 2.3;a mount flexibly supporting the gear system through the static structure of the engine and configured to accommodate movement between the gear system and the static structure; anda lubrication system configured to provide lubricant to the gear system and remove thermal energy from the gear system; andwherein the plurality of fan blades is less than 26 and the plurality of ...

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Номер записи: 19
08-12-2016 дата публикации

GEARED ARCHITECTURE FOR A GAS TURBINE ENGINE

Номер: US20160356225A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gear system for a gas turbine engine includes a planet gear system which includes an output attached to a carrier for rotating a first fan assembly in a first direction. A star gear system includes an output attached to a ring gear for rotating a second fan assembly in a second direction. A sun gear of the star gear system is mechanically attached to a sun gear of the planet gear system.

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Номер записи: 20
20-04-2017 дата публикации

FIXED SUPPORT AND OIL COLLECTOR SYSTEM FOR RING GEAR

Номер: US20170108110A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

According to various embodiments, disclosed is a fixed support and gutter system for a ring gear of a turbine engine gear train, the fixed support and gutter system comprising a support structure having an undulating path which forms at least a portion of a gutter structure for capturing oil shed off the gear train. 1. A fixed support and gutter system for a gear train of a turbine engine , the fixed support and gutter system comprising:a support structure coupled between at least one ring gear of the gear train and an engine case of the turbine engine; and 'wherein the support structure comprises an undulating path forming at least a portion of the gutter structure integrated with the support structure, and wherein the gutter structure is configured to capture oil shed from the gear train.', 'a gutter structure;'}2. The fixed support and gutter system of claim 1 , the gutter structure comprising at least one discharge passage coupled to the gutter structure for conducting captured oil.3. The fixed support and gutter system of claim 1 , wherein the gutter structure is positioned circumferentially around the gear train.4. The fixed support and gutter system of claim 1 , wherein the gutter structure comprises a forward oil collector positioned forward of the gear train claim 1 , and an aft oil collector claim 1 , which is positioned aft of the gear train.5. The fixed support and gutter system of claim 4 , further comprising a fluid passage between the forward oil collector and the aft oil collector.6. The fixed support and gutter system of claim 4 , wherein at least one of the forward oil collector and the aft oil collector is formed from the undulating path of the support structure.7. The fixed support and gutter system of claim 6 , the undulating path of the support structure forming one of the forward oil collector or the aft oil collector claim 6 , wherein the other of the forward oil collector or aft oil collector does not form part of the support structure.8. ...

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Номер записи: 21
13-10-2016 дата публикации

LUBRICANT CIRCULATION SYSTEM AND METHOD OF CIRCULATING LUBRICANT IN A GAS TURBINE ENGINE

Номер: US20160298543A1
Автор: Gabriel L. Suciu, William G. Sheridan, SUCIU GABRIEL L, SHERIDAN WILLIAM G, Suciu Gabriel L., Sheridan William G.
Принадлежит:

A lubricant circulation system and method of circulating lubricant in a gas turbine engine are disclosed. The lubricant circulation system includes a nose cone having an aperture communicating air to an interior space of the nose cone, a heat exchanger disposed in the interior space, and a lubricant circulation pathway contained within a forward portion of the gas turbine engine and configured to circulate lubricant through the heat exchanger.

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Номер записи: 22
29-12-2016 дата публикации

GEARED TURBOFAN WITH INDEPENDENT FLEXIBLE RING GEARS AND OIL COLLECTORS

Номер: US20160376984A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A geared turbofan engine includes a fan rotatable about an engine axis. A compressor section compresses air and delivers the compressed air to a combustor where the compressed air is mixed with fuel and ignited to drive a turbine section that in turn drives the fan and the compressor section. A gear system is driven by the turbine section for driving the fan at a speed different than the turbine section. The gear system includes a carrier attached to a fan shaft. A plurality of planet gears are supported within the carrier. Each of the plurality of planet gears includes a first row of gear teeth and a second row of gear teeth supported within the carrier. A sun gear is driven by a turbine section. The sun gear is in driving engagement with the plurality of planet gears. At least two separate ring gears circumscribe the plurality of planet gears. Each of the at least two ring gears are supported by a respective flexible ring gear mount that enables movement relative to an engine static structure ...

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Номер записи: 23
19-09-2017 дата публикации

Lubrication systems for gearbox assemblies

Номер: US9765875B2
Автор: SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: SIKORSKY AIRCRAFT CORP, Sikorsky Aircraft Corporation

A gearbox assembly includes a housing with a housing interior. A sump is disposed within a lower region of the gearbox housing. A lubricated transmission element is arranged in the housing interior above the sump. A lubricant impoundment is arranged within the housing and in series between the transmission element and the sump such that lubricant flowing in a primary lubricant flow path between the transmission element and the sump is impounded in the lubricant impoundment, thereby providing a supply of lubricant for a secondary lubricant flow path disposed within the gearbox housing.

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Номер записи: 24
08-06-2017 дата публикации

GEAR DRIVEN GAS TURBINE ENGINE ASSEMBLY

Номер: US20170159798A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gas turbine engine includes a planetary gear system including a sun gear, intermediate gears, and a ring gear. A lubricant recovery system for the planetary gear system includes fluid passages that extend through the planetary gear system. A gutter is located radially outward from the planetary gear system for collecting lubricant. At least a portion of the gutter is rigidly attached to the ring gear.

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Номер записи: 25
29-08-2017 дата публикации

Reduced misalignment gear system

Номер: US0009745862B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A lubrication system for a fan drive planetary gear system according to an exemplary aspect of the present disclosure includes, among other things, a stationary first bearing configured to receive a lubricant from a lubricant input, the stationary first bearing is located adjacent a fan drive shaft. A second bearing is configured to rotate with the fan drive shaft, the first bearing engages the second bearing and is configured to transfer the lubricant from the first bearing to the second bearing and into at least one fluid passage in the fan drive shaft. A conduit fluidly connects the at least one passage in the fan drive shaft with at least one component on the fan drive gear system.

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Номер записи: 26
14-09-2017 дата публикации

GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20170260911A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gas turbine engine includes a gear system that transfers power input from a fan drive turbine to a fan through a gear system with a gear reduction ratio greater than 2.3. A flexible mount supports the gear system and accommodates movement of the gear system relative to an engine static structure. A lubrication system provides lubricant flow to a plurality of rotating components including bearing systems and the gear system and has a maximum capacity for removing thermal energy less than or equal to no more than 2% of power input into the gear system.

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Номер записи: 27
15-08-2017 дата публикации

Fan drive gear system integrated carrier and torque frame

Номер: US0009732839B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A method of assembling a fan drive gear system includes the steps of installing spherical bearings into respective races to provide a plurality of bearing assemblies, mounting at least one of the bearing assemblies onto a corresponding shaft of a torque frame, each of the shafts fixed relative to one another, installing at least one gear onto at least one of the bearing assemblies, the gears meshing with a ring gear and a centrally located sun gear and grounding the torque frame to a static structure to prevent rotation of the torque frame.

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Номер записи: 28
08-12-2016 дата публикации

TURBINE ENGINE WITH A TURBO-COMPRESSOR

Номер: US20160356218A1
Автор: Daniel B. Kupratis, William G. Sheridan, KUPRATIS DANIEL B, SHERIDAN WILLIAM G, Kupratis Daniel B., Sheridan William G.
Принадлежит:

A turbine engine is provided that includes a first rotor, a second rotor and a combustor section. The first rotor includes a first set of compressor blades. The second rotor is adjacent the first rotor. The second rotor includes a second set of compressor blades and a set of turbine blades respectively connected to the second set of compressor blades. The combustor section is configured to receive air compressed by the first and the second sets of compressor blades. The combustor section is also configured to provide combustion products to the set of turbine blades.

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Номер записи: 29
15-08-2017 дата публикации

Turbofan engine main bearing arrangement

Номер: US0009732629B2
Автор: William G. Sheridan, Frederick M. Schwarz, SHERIDAN WILLIAM G, SCHWARZ FREDERICK M, Sheridan William G., Schwarz Frederick M.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A turbofan engine (20; 300; 400) comprises a fan (28), a fan drive gear system (60), a fan shaft (120) coupling the fan drive gear system to the fan, a low spool, an intermediate spool, and a core spool. The low spool engages at least three main bearings of which at least two are non-thrust bearings and at least one is a thrust bearing. The fan shaft engages at least two bearings (148, 150). The core spool engages at least two bearings (250, 260). The intermediate spool engages at least two of said bearings (220, 200, 230; 220, 200, 230-2; 200, 220, 230-3).

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Номер записи: 30
03-10-2017 дата публикации

Oil baffle for gas turbine fan drive gear system

Номер: US0009777825B2
Автор: William G. Sheridan, Michael E. McCune, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, Sheridan William G., McCune Michael E.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A turbine engine includes a housing supporting compressor and turbine sections. An epicyclic gear train includes a carrier, a sun gear and intermediate gears arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle is supported relative to the carrier and includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on at least one of the sun gear and the intermediate gears. A spray bar is external to the carrier and is in communication with the lubrication passage. The spray bar terminates near the sun gear for directing lubricant on the sun gear.

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Номер записи: 31
29-12-2016 дата публикации

WINDMILL AND NEGATIVE-G OIL SYSTEM FOR GEARED TURBOFAN ENGINES

Номер: US20160376988A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A lubrication system for a gear system of a geared turbofan engine, the lubrication system includes an upper strut including an upper sump. A lower strut includes a lower sump. The upper strut and the lower strut support a gear system. A conduit extends from the upper sump within the upper strut to the lower sump within the lower strut. A lubricant collector receives lubricant exhausted from the gear system and directing the received lubricant into one of the lower sump and the upper sump. A pump in communication with the conduit for drawing lubricant from at least one of the lower sump and the upper sump and communicates lubricant to a lubricant inlet of the gear system. The pump draws lubricant from the lower sump when operating within a positive g-force environment and draws lubricant from the upper sump when operating within a negative g-force environment. A geared turbofan engine is also disclosed.

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Номер записи: 32
30-01-2018 дата публикации

Turbo-compressor with geared turbofan

Номер: US0009879694B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP

A gas turbine engine includes a fan section. A turbo-compressor section is connected to the fan section through a speed change mechanism.

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Номер записи: 33
11-08-2016 дата публикации

GEAR REDUCTION FOR LOWER THRUST GEARED TURBOFAN

Номер: US20160230674A1
Автор: Frederick M. Schwarz, William G. Sheridan, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, Schwarz Frederick M., Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan rotor having a hub and a plurality of fan blades extending radially outwardly of the hub. A compressor is positioned downstream of the fan rotor, and has a first compressor blade row defined along a rotational axis of the fan rotor and the compressor rotor. A gear reduction is positioned axially between the first compressor blade row and the fan rotor, and includes a ring gear and a carrier. The carrier has an axial length and the ring gear has an outer diameter. A ratio of the axial length to the outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40. The gear reduction is connected to drive the hub to rotate. A method of designing a gas turbine engine is also disclosed.

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Номер записи: 34
14-11-2017 дата публикации

Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

Номер: US0009816443B2
Автор: William G. Sheridan, Karl L. Hasel, SHERIDAN WILLIAM G, HASEL KARL L, Sheridan William G., Hasel Karl L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram ° R)/(518.7 ° R)]0.5, where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio. A low and high pressure turbine in communication with a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages.

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Номер записи: 35
23-01-2018 дата публикации

Oil baffle for gas turbine fan drive gear system

Номер: US0009874274B2
Автор: William G. Sheridan, Michael E. McCune, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, Sheridan William G., McCune Michael E.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A geared architecture includes a carrier that has spaced apart walls with circumferentially spaced mounts that interconnect the walls. The mounts provide circumferentially spaced apart apertures between the mounts at an outer circumference of the carrier. A sun gear and intermediate gears are arranged about and intermesh with the sun gear. The intermediate gears are supported by the carrier. The intermediate gears extend through the apertures to intermesh with a ring gear. A baffle is supported relative to the carrier and includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on the at least one of the sun gear and intermediate gears. The lubrication passage includes a primary passage that extends laterally between the walls and first and second passages and is in communication with the primary passage and respectively terminates near the sun gear and intermediate gears.

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Номер записи: 36
04-08-2016 дата публикации

TURBO-COMPRESSOR WITH GEARED TURBOFAN

Номер: US20160222814A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section. A turbo-compressor section is connected to the fan section through a speed change mechanism.

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Номер записи: 37
09-01-2018 дата публикации

Flexible coupling for geared turbine engine

Номер: US0009863326B2
Автор: William G. Sheridan, John R. Otto, SHERIDAN WILLIAM G, OTTO JOHN R, Sheridan William G., Otto John R.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

A gas turbine engine includes a fan shaft arranged along an engine central axis, a frame supporting the fan shaft, a gear system rotatably coupled with the fan shaft, and a flexible coupling at least partially supporting the gear system. The flexible coupling defines, with respect to the engine central axis, a torsional stiffness TS and a lateral stiffness LS such that a ratio of TS/LS is greater than or equal to about 2.0 to reduce loads on the gear system from misalignment of the gear system with respect to the engine central axis.

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Номер записи: 38
27-07-2017 дата публикации

GEARED GAS TURBINE ENGINE

Номер: US20170211484A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A turboshaft engine includes a high speed spool that connects a high pressure compressor with a high pressure turbine. A low speed spool connects a low pressure compressor with a low pressure turbine. A speed change mechanism includes an input that is in communication with the low spool and a fixed gear ratio. An output turboshaft is in communication with an output of the speed change mechanism.

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Номер записи: 39
10-11-2016 дата публикации

LIGHTWEIGHT JOURNAL SUPPORT PIN

Номер: US20160326902A1
Автор: Michael E. McCune, William G. Sheridan, William P. Ogden, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, OGDEN WILLIAM P, McCune Michael E., Sheridan William G., Ogden William P.
Принадлежит:

A journal support pin to support intermediate gears for use in gas turbine engine comprises a titanium body, and an outer surface outside of the titanium body having a surface hardness that is harder than the body. A gas turbine engine and a method of forming a journal support pin to support intermediate gears for use in gas turbine engine are also disclosed.

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Номер записи: 40
30-01-2018 дата публикации

Hybrid drive for gas turbine engine

Номер: US0009878796B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A gas turbine engine has a fan drive turbine for selectively driving a fan rotor. A drive shaft between the fan drive turbine and the fan rotor includes a clutch, and an electric motor. The electric motor is positioned such that it is not downstream of a flow path relative to the fan drive turbine. A method of operating a gas turbine engine is also disclosed.

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Номер записи: 41
18-08-2016 дата публикации

TURBINE ENGINE WITH A TURBO-COMPRESSOR

Номер: US20160237895A1
Автор: Daniel B. Kupratis, William G. Sheridan, KUPRATIS DANIEL B, SHERIDAN WILLIAM G, Kupratis Daniel B., Sheridan William G.
Принадлежит:

A turbine engine is provided that includes a turbo-compressor and a combustor section. The turbo-compressor includes a compressor section and a turbine section. The combustor section is fluidly coupled between the compressor section and the turbine section. The compressor section includes a first number of stages. The turbine section includes a second number of stages that is different than the first number.

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Номер записи: 42
06-10-2016 дата публикации

AUXILIARY OIL SYSTEM FOR NEGATIVE GRAVITY EVENT

Номер: US20160290229A1
Автор: William G. Sheridan, Paul H. Dolman, SHERIDAN WILLIAM G, DOLMAN PAUL H, Sheridan William G., Dolman Paul H.
Принадлежит:

A method of supplying a bearing compartment with fluid includes the steps of pumping a fluid from a main reservoir to a bearing compartment through a main supply passage which includes one or more passage segments. During a positive gravity condition, the pumping step is performed using a main pump arranged in the main supply passage. The main reservoir includes upper and lower portions. The main supply passage is in fluid communication with the lower portion. Fluid is provided from the main reservoir to the bearing compartment through a secondary supply passage that is fluidly connected to at least one segment of the main supply passage in response to a negative gravity condition. The secondary supply passage is in fluid communication with the upper portion.

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Номер записи: 43
25-08-2016 дата публикации

IN-LINE DEAERATOR DEVICE FOR WINDMILL-AUXILIARY OIL SYSTEM FOR FAN DRIVE GEAR SYSTEM

Номер: US20160245117A1
Автор: Francis Parnin, William G. Sheridan, John S. Groholski, PARNIN FRANCIS, SHERIDAN WILLIAM G, GROHOLSKI JOHN S, Parnin Francis, Sheridan William G., Groholski John S.
Принадлежит:

A lubrication system for a fan drive gear system of a turbofan engine includes an auxiliary pump for communicating lubricant to bearings of a gear system. A deaerator is disposed between a lubricant source and the auxiliary pump for separating gases is contained within the lubricant. The deaerator includes a vane section for inducing a radial flow in the lubricant for separating gas from the lubricant. A gas turbine engine and method are also disclosed.

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Номер записи: 44
06-02-2018 дата публикации

Turbofan engine bearing and gearbox arrangement

Номер: US0009885282B2
Автор: William G. Sheridan, Michael E. McCune, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, Sheridan William G., McCune Michael E.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A turbofan engine (20) has a fan shaft (120) coupling a fan drive gear system (60) to the fan (28). A low spool comprises a low pressure turbine (50) and a low shaft (56) coupling the low pressure turbine to the fan drive gear system. A core spool comprises a high pressure turbine (46), a compressor (44), and a core shaft (52) coupling the high pressure turbine to the core spool compressor. A first bearing (150) engages the fan shaft, the first bearing being a thrust bearing. A second bearing (160) engages the fan shaft on an opposite side of the fan drive gear system from the first bearing, the second bearing being a roller bearing. A third bearing (180) engages the low spool shaft and the fan shaft.

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Номер записи: 45
06-04-2017 дата публикации

GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20170096944A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an axis;a fan pressure ratio across a fan blade alone of less than 1.45;a bypass duct;a compressor section;{'b': '10', 'a bypass ratio greater than , the 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;'}a combustor in fluid communication with the compressor section;{'b': '5', 'a fan drive turbine in communication with the combustor, the fan drive turbine comprising a pressure ratio greater than about , wherein the fan drive turbine further includes an inlet having an inlet pressure, and an outlet that is prior to any exhaust nozzle and having an outlet pressure, and the pressure ratio of the fan drive turbine is a ratio of the inlet pressure to the outlet pressure;'}a gear system configured to provide a speed reduction between the fan drive turbine and the fan and transfer power input from the fan drive turbine to the fan at an efficiency greater than 98%, wherein the gear system further comprises a gear reduction ratio of greater than 2.3; anda lubrication system configured to provide lubricant to the gear system and remove thermal energy from the gear system; andwherein the plurality of fan blades is less than twenty (20) and the plurality of fan drive turbine rotors is less than six (6) fan drive turbine rotors, and a ratio between the number of fan blades and the number of fan drive turbine ...

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Номер записи: 46
10-11-2016 дата публикации

LUBRICATION SYSTEM FOR GEARED GAS TURBINE ENGINE

Номер: US20160326906A1
Автор: William G. Sheridan, Gabriel L. Suciu, SHERIDAN WILLIAM G, SUCIU GABRIEL L, Sheridan William G., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An oil cooling system is provided. The system may comprise an oil inlet, a manifold fluidly coupled to the oil inlet, and a support strut comprising internal tubing fluidly coupled to the manifold. A heat exchanger may be fluidly coupled to the internal tubing of the support strut. A nose cone may be disposed forward of the heat exchanger and configured to rotate about an axis. The heat exchanger may be radially inward from a portion of the nose cone. A gas turbine engine is also provided. The gas turbine engine may comprise an epicyclic gear system and fan mechanically coupled to the epicyclic gear system. The fan may be configured to rotate about an axis. A nose cone may be coupled to the fan and configured to rotate about the axis. A heat exchanger may be aft of the fan and in fluid communication with the epicyclic gear system.

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Номер записи: 47
12-12-2017 дата публикации

Gear system architecture for gas turbine engine

Номер: US0009840969B2
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system.

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Номер записи: 48
03-10-2017 дата публикации

Lubrication system for geared gas turbine engine

Номер: US0009777595B2
Автор: William G. Sheridan, Gabriel L. Suciu, SHERIDAN WILLIAM G, SUCIU GABRIEL L, Sheridan William G., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION, UNITED TECHNOLOGIES CORP, United Technologies Corporation

An oil cooling system is provided. The system may comprise an oil inlet, a manifold fluidly coupled to the oil inlet, and a support strut comprising internal tubing fluidly coupled to the manifold. A heat exchanger may be fluidly coupled to the internal tubing of the support strut. A nose cone may be disposed forward of the heat exchanger and configured to rotate about an axis. The heat exchanger may be radially inward from a portion of the nose cone. A gas turbine engine is also provided. The gas turbine engine may comprise an epicyclic gear system and fan mechanically coupled to the epicyclic gear system. The fan may be configured to rotate about an axis. A nose cone may be coupled to the fan and configured to rotate about the axis. A heat exchanger may be aft of the fan and in fluid communication with the epicyclic gear system.

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Номер записи: 49
08-12-2016 дата публикации

TURBINE ENGINE WITH A TURBO-COMPRESSOR

Номер: US20160356244A1
Автор: Daniel B. Kupratis, William G. Sheridan, KUPRATIS DANIEL B, SHERIDAN WILLIAM G, Kupratis Daniel B., Sheridan William G.
Принадлежит:

A turbine engine is provided that include a turbo-compressor, a combustor section and a nacelle which houses the turbo-compressor and the combustor section. The turbo-compressor includes a compressor section and a turbine section. The combustor section is fluidly coupled between the compressor section and the turbine section. The nacelle includes a thrust reverser.

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Номер записи: 50
13-10-2016 дата публикации

SPEED SENSOR FOR A GAS TURBINE ENGINE

Номер: US20160298485A1
Автор: William G. Sheridan, Jason Husband, SHERIDAN WILLIAM G, HUSBAND JASON, Sheridan William G., Husband Jason
Принадлежит:

A gas turbine engine includes a speed change mechanism. A fan drive shaft has a radially extending surface that is attached to the speed change mechanism. A speed sensor is located adjacent the radially extending surface.

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Номер записи: 51
11-10-2016 дата публикации

Gear carrier flex mount lubrication

Номер: US0009464708B2
Автор: James B. Coffin, Michael E. McCune, William G. Sheridan, COFFIN JAMES B, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, Coffin James B., McCune Michael E., Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

An exemplary method of lubricating a turbomachine interface includes, among other things, securing a carrier relative to a torque frame using a flexure pin, and lubricating an interface of the flexure pin using a lubricant that has moved through a lubricant passage in the carrier and a lubricant passage in the flexure pin.

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Номер записи: 52
29-09-2016 дата публикации

METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE

Номер: US20160281610A1
Автор: William G. Sheridan, Karl L. Hasel, SHERIDAN WILLIAM G, HASEL KARL L, Sheridan William G., Hasel Karl L.
Принадлежит:

A gas turbine engine includes an engine centerline longitudinal axis and a fan section including a fan with fan blades and rotatable about the engine centerline longitudinal axis. A low corrected fan tip speed less than about 1400 ft/sec and the low corrected fan tip speed is an actual fan tip speed determined at an ambient temperature divided by [(Tram° R)/(518.7° R)]0.5, where T represents the ambient temperature in degrees Rankine. A bypass ratio greater than about 11 and a speed reduction device having a gear system with a gear ratio and a plurality of bearing systems. A low and high speed spool including a low and high pressure turbine and a first and second shaft, respectively. The first and second shafts are concentric and mounted via at least one of the bearing systems for rotation about the engine centerline longitudinal axis and the first shaft is in communication with the fan through the speed reduction device and the low pressure turbine includes four stages.

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Номер записи: 53
04-08-2016 дата публикации

DUAL DIRECTION WINDMILL PUMP FOR GEARED TURBOFAN ENGINE

Номер: US20160222975A1
Автор: William G. Sheridan, Reade W. James, Jacob Peter Mastro, SHERIDAN WILLIAM G, JAMES READE W, MASTRO JACOB PETER, Sheridan William G., James Reade W., Mastro Jacob Peter
Принадлежит:

A lubrication system includes a shaft rotatable about an axis, a lubrication pump configured to supply a lubricant flow to a gear system, and a gear train coupled to the shaft and configured to drive the lubrication pump in a first direction responsive to rotation of the shaft in both the first direction and a second direction. A gas turbine engine and method are also disclosed.

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Номер записи: 54
11-05-2017 дата публикации

GEARED TURBOMACHINE FAN AND COMPRESSOR ROTATION

Номер: US20170130673A1
Автор: William G. Sheridan, Michael E. McCune, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, Sheridan William G., McCune Michael E.
Принадлежит: United Technologies Corp

An exemplary gas turbine engine includes a fan section including a fan rotor and at least one fan blade. A fan pressure ratio across the at least one fan blade is less than 1.45, noninclusive of the pressure across any fan exit guide vane system. The engine further includes a low-pressure compressor having a low-pressure compressor rotor that rotates together with the fan rotor at a common speed in operation, and a geared architecture that drives the low-pressure compressor rotor and the fan rotor. The geared architecture has a gear reduction ratio of greater than 2.5. The engine further includes a high-pressure compressor having a pressure ratio greater than 20, a low-pressure turbine having a pressure ratio greater than 5, and a bypass ratio greater than 10.

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Номер записи: 55
27-03-2018 дата публикации

Compound star gear system with rolling element bearings

Номер: US9926850B2
Автор: SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORP, United Technologies Corporation

A compound star gear system includes a sun gear rotatable about a first axis that drives a first plurality of star gears rotatable about a plurality of fixed axes. The first plurality of star gears drives a second plurality of star gears spaced axially apart from the first plurality of star gears. The second plurality of star gears drive the ring gear that in turn drives a fan drive shaft.

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Номер записи: 56
22-11-2016 дата публикации

Geared turbofan engine gearbox arrangement

Номер: US0009500126B2
Автор: Frederick M. Schwarz, William G. Sheridan, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, Schwarz Frederick M., Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A three-spool turbofan engine (20) has a variable fan nozzle (35). The fan blades have a peak tip radius RT and an inboard leading edge radius RH at an inboard boundary of the flowpath. A ratio of RH to RT is less than about 0.40.

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Номер записи: 57
29-12-2016 дата публикации

ROLLING ELEMENT CAGE FOR GEARED TURBOFAN

Номер: US20160377167A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gear system for a turbofan engine assembly includes a sun gear rotatable about an engine centerline, a non-rotatable ring gear, and a rotating carrier that drives a fan. A plurality of planet gears is intermeshed between the sun gear and the ring gear. A rolling element bearing assembly supports rotation of the planet gear on the carrier. The rolling element bearing assembly includes a rolling element between an inner race and an outer race separated by a cage. A first passage for lubricant through the planet gear. A second passage is in communication with the first passage for communicating lubricant through the inner race to an interface between the inner race and the cage. A geared turbofan engine is also disclosed.

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Номер записи: 58
06-03-2018 дата публикации

Gas turbine engine with distributed fans with drive control

Номер: US0009909495B2
Автор: Gabriel L. Suciu, Wesley K. Lord, Jesse M. Chandler, Michael Winter, Michael E. McCune, William G. Sheridan, SUCIU GABRIEL L, LORD WESLEY K, CHANDLER JESSE M, WINTER MICHAEL, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, Suciu Gabriel L., Lord Wesley K., Chandler Jesse M., Winter Michael, McCune Michael E., Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A gas turbine engine comprises a plurality of fan rotors. A gas generator comprises at least one compressor rotor, at least one gas generator turbine rotor, a combustion section, and a fan drive turbine downstream of at least one gas generator turbine rotor. A shaft is configured to be driven by the fan drive turbine. The shaft engages gears to drive the plurality of fan rotors. A system controls the amount of power supplied to the plurality of fan rotors. A method of operating a gas turbine engine is also disclosed.

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Номер записи: 59
01-06-2017 дата публикации

GEARED TURBOFAN WITH FOUR STAR/PLANETARY GEAR REDUCTION

Номер: US20170152756A1
Автор: Frederick M. Schwarz, William G. Sheridan, Michael E. McCune, Gabriel L. Suciu, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, MCCUNE MICHAEL E, SUCIU GABRIEL L, Schwarz Frederick M., Sheridan William G., McCune Michael E., Suciu Gabriel L.
Принадлежит:

A gas turbine engine has a fan rotor, a turbine rotor driving the fan rotor, and an epicyclic gear reduction positioned between the fan rotor and the turbine rotor. The epicyclic gear reduction includes a ring gear, a sun gear, and no more than four intermediate gears that engage the sun gear and the ring gear. The fan drive turbine is configured to drive the sun gear to, in turn, drive the ring gears to, in turn, drive the fan rotor.

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Номер записи: 60
25-08-2016 дата публикации

GEARED TURBINE ENGINE

Номер: US20160245184A1
Автор: Frederick M. Schwarz, William G. Sheridan, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbine engine is provided that includes a fan rotor, a first compressor rotor, a second compressor rotor, a third compressor rotor, a first turbine rotor, a second turbine rotor, a third turbine rotor and a gear train. The fan rotor and the first compressor rotor are connected to the first turbine rotor through the gear train. The second compressor rotor is connected to the second turbine rotor. The third compressor rotor is connected to the third turbine rotor.

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Номер записи: 61
22-12-2016 дата публикации

LUBRICATION SYSTEMS FOR GEARBOX ASSEMBLIES

Номер: US20160369887A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A gearbox assembly includes a housing with a housing interior. A sump is disposed within a lower region of the gearbox housing. A lubricated transmission element is arranged in the housing interior above the sump. A lubricant impoundment is arranged within the housing and in series between the transmission element and the sump such that lubricant flowing in a primary lubricant flow path between the transmission element and the sump is impounded in the lubricant impoundment, thereby providing a supply of lubricant for a secondary lubricant flow path disposed within the gearbox housing.

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Номер записи: 62
23-01-2018 дата публикации

Method of assembly for gas turbine fan drive gear system

Номер: US0009874150B2
Автор: Michael E. McCune, William G. Sheridan, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, McCune Michael E., Sheridan William G.
Принадлежит: United Technologies Corporation, UNITED TECHNOLOGIES CORP

A method of assembling an epicyclic gear train comprises the steps of providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining spaced apart apertures provided at an outer circumference of the carrier. Gear pockets are provided between the walls and extend to the apertures. A central opening is in at least one of the walls. A plurality of intermediate gears are inserted through the central opening and move the intermediate gears radially outwardly into the gear pockets to extend into the apertures. A sun gear is inserted through the central opening. The plurality of intermediate gears is moved radially inwardly to engage the sun gear.

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Номер записи: 63
28-09-2017 дата публикации

PLANETARY GEAR SYSTEM ARRANGEMENT WITH AUXILIARY OIL SYSTEM

Номер: US20170276046A1
Автор: Michael E. McCune, William G. Sheridan, Lawrence E. Portlock, MCCUNE MICHAEL E, SHERIDAN WILLIAM G, PORTLOCK LAWRENCE E, McCune Michael E., Sheridan William G., Portlock Lawrence E.
Принадлежит:

A method of designing a gas turbine engine includes configuring a speed reduction device for driving a fan and configuring a lubrication system for lubricating components across a rotation gap. The lubrication system includes a lubricant input. A stationary first bearing receives lubricant from the lubricant input and has a first race in which lubricant flows. A second bearing for rotation is within the first bearing. The second bearing has a first opening in registration with the first race such that lubricant may flow from the first race through the first opening into a first conduit. The first bearing is configured to also include a second race into which lubricant flows. The second bearing has a second opening in registration with the second race such that lubricant may flow from the second race through the second opening into a second conduit. The first and second conduits deliver lubricant to distinct locations.

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Номер записи: 64
18-05-2017 дата публикации

MONITORING SYSTEM FOR NON-FERROUS METAL PARTICLES

Номер: US20170138217A1
Автор: Frederick M. Schwarz, William G. Sheridan, SCHWARZ FREDERICK M, SHERIDAN WILLIAM G, Schwarz Frederick M., Sheridan William G.
Принадлежит:

According to one aspect of the present disclosure, a debris monitoring system is disclosed that includes a fan, a geared architecture operatively coupled to the fan. The geared architecture includes a component having a non-ferrous metal coating. A scavenge pump is in fluid communication with the geared architecture via a lubrication sump. A non-ferrous chip detector is situated downstream of the geared architecture, but upstream of the scavenge pump. A controller is configured to determine a lubrication condition of the component based on a signal received from the non-ferrous chip detector, and command a status indicator in response thereto. 1. A debris monitoring system , comprising:a fan;a geared architecture operatively coupled to the fan, and comprising a component having a non-ferrous metal coating;a scavenge pump in fluid communication with the geared architecture via a lubrication sump;a non-ferrous chip detector situated downstream of the geared architecture, but upstream of the scavenge pump; anda controller configured to determine a lubrication condition of the component based on a signal received from the non-ferrous chip detector, and command a status indicator in response thereto.2. The debris monitoring system of claim 1 , wherein the component is a journal pin.3. The debris monitoring system of claim 1 , wherein the non-ferrous metal coating comprises one or more of copper claim 1 , silver claim 1 , and lead.4. The debris monitoring system of claim 1 , wherein the non-ferrous chip detector is situated downstream of the lubrication sump.5. The debris monitoring system of claim 4 , wherein the scavenge pump is operative to pump lubricant from the lubrication sump claim 4 , and from one or more additional lubrication sumps that are located in parallel flow paths that omit the geared architecture.6. The debris monitoring system of claim 1 , wherein the non-ferrous chip detector at least partially surrounds a portion of conduit that carries lubricant ...

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Номер записи: 65
05-01-2017 дата публикации

AUXILIARY OIL SYSTEM FOR GEARED GAS TURBINE ENGINE

Номер: US20170002738A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine comprises a fan drive turbine, a fan rotor, and a gear reduction driven by the fan drive turbine to, in turn, drive the gear architecture. A main oil supply system supplies oil to components within the gear reduction, and an auxiliary oil supply system. The auxiliary oil system operates to ensure that the gear reduction will be adequately supplied with lubricant for at least seconds at power should the main oil supply system fail. 1. A gas turbine engine comprising:a fan drive turbine, a fan rotor, and a gear reduction driven by said fan drive turbine to, in turn, drive said gear architecture, a main oil supply system for supplying oil to components within said gear reduction, and an auxiliary oil supply system; andsaid auxiliary oil system being operable to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail.2. The gas turbine engine as set forth in claim 1 , wherein said gear reduction includes a sun gear being driven by said fan drive turbine to drive intermediate gears that engage a ring gear.3. The gas turbine engine as set forth in claim 2 , wherein said sun gear claim 2 , said intermediate gears and said ring gear are enclosed in a bearing compartment claim 2 , which captures oil removed via a scavenge line connected to a main oil pump.4. The gas turbine engine as set forth in claim 3 , wherein said main oil pump has a gutter that directs scavenged oil to a main oil tank.5. The gas turbine engine as set forth in claim 4 , wherein oil in said main oil tank feeds a main pump pressure stage which then delivers oil to said gear reduction.6. The gas turbine engine as set forth in claim 5 , wherein oil from said main pump pressure stage passes through a lubrication system that includes at least one filter and at least one heat exchanger to cool the oil.7. The gas turbine engine as set forth in claim 4 , wherein said gear reduction is surrounded by an ...

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Номер записи: 66
27-10-2016 дата публикации

GEARED TURBOFAN WITH IMPROVED GEAR SYSTEM MAINTAINABILITY

Номер: US20160312696A1
Автор: William G. Sheridan, SHERIDAN WILLIAM G, Sheridan William G.
Принадлежит:

A disclosed gas turbine engine includes a fan section including a hub supporting a plurality of fan blades rotatable about an axis, and a bearing assembly supporting rotation of the hub about the axis. A compressor section is in fluid communication with a combustor and a turbine is in fluid communication with the compressor section. A speed reduction device driven by the turbine section for rotating the fan about the axis is mounted forward of the bearing assembly supporting rotation of the hub.

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Номер записи: 67
03-11-2016 дата публикации

FAN DRIVE GEAR SYSTEM INCLUDING A TWO-PIECE FAN SHAFT WITH LUBRICANT TRANSFER LEAKAGE RECAPTURE

Номер: US20160319830A1
Автор: William G. Sheridan, Karl D. Blume, SHERIDAN WILLIAM G, BLUME KARL D, Sheridan William G., Blume Karl D.
Принадлежит:

A disclosed fan drive gear system for a gas turbine engine includes a first fan shaft coupled to a second fan shaft, a first shaft support bearing assembly disposed about the first fan shaft and a second shaft support bearing assembly disposed about the second fan shaft. A planetary gear system is coupled to the second fan shaft. A transfer bearing is configured to receive lubricant from a lubricant input and is positioned between the first and second fan shaft support bearings. A second bearing is configured to rotate with the second fan shaft and receive lubricant from the transfer bearing and communicate lubricant to at least one lubricant passage and a conduit fluidly connecting the at least one lubricant passage to the planetary gear system.

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Номер записи: 68
12-04-2012 дата публикации

Planetary gear system arrangement with auxiliary oil system

Номер: US20120088624A1
Автор: William G. Sheridan
Принадлежит: Individual

A lubricating system for lubricating components across a rotation gap has an input that provides lubricant to a stationary first bearing. The first bearing has an inner first race in which lubricant flows. A second bearing rotates within the first bearing and has a first opening in registration with the inner first race such that lubricant may flow from the inner first race through the first opening into a first conduit.

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Номер записи: 69
02-08-2012 дата публикации

PLANETARY GEAR SYSTEM ARRANGEMENT WITH AUXILIARY OIL SYSTEM

Номер: US20120192570A1
Автор: McCune Michael E., Portlock Lawrence E., Sheridan William G.
Принадлежит:

A gas turbine engine has a fan, first and second compressor stages, first and second turbine stages. The first turbine stage drives the second compressor stage as a high spool. The second turbine stage drives the first compressor stage as part of a low spool. A gear train drives the fan with the low spool, such that the fan and first compressor stage rotate in the same direction. The high spool operates at higher pressures than the low spool. A lubrication system is also disclosed. 1. A gas turbine engine comprising:a fan, a first compressor stage and a second compressor stage;a first turbine stage and a second turbine stage, and wherein said first turbine stage drives said second compressor stage as a high spool, and wherein said second turbine stage drives said first compressor stage as part of a low spool; anda gear train driving said fan with said low spool, and such that said fan and said first compressor stage rotate in the same direction, and wherein said high spool operates at higher pressures than said low spool.2. The gas turbine engine of claim 1 , further comprising said gear train having a planetary gear claim 1 , a sun gear claim 1 , a stationary ring gear claim 1 , a carrier in which said planetary gear is mounted claim 1 , and said carrier mounted for rotation about said sun gear and driving said fan.3. The gas turbine engine of claim 2 , wherein a lubricating system is provided for said gear train.4. The gas turbine engine of claim 3 , wherein the lubricating system includes a lubricant input claim 3 , there being a stationary first bearing receiving lubricant from said lubricant input claim 3 , said first bearing having an inner first race in which lubricant flows claim 3 , and a second bearing for rotation within said first bearing claim 3 , said second bearing having a first opening in registration with said inner first race such that lubricant may flow from said inner first race through said first opening into a first conduit.5. The gas turbine ...

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Номер записи: 70
01-11-2012 дата публикации

FAN DRIVE GEAR SYSTEM INTEGRATED CARRIER AND TORQUE FRAME

Номер: US20120272762A1
Автор: Sheridan William G.
Принадлежит:

A fan gear drive system includes a torque frame having a base with integrated gear shafts circumferentially spaced relative to one another. Each shaft provides a shaft axis. A bearing assembly is mounted on each of the gear shafts and provides a bearing assembly. The bearing assembly includes a spherical bearing configured to permit angular movement of the bearing axis relative to the shaft axis. 1. A fan gear drive system comprising:a torque frame comprising a base with integrated gear shafts circumferentially spaced relative to one another, each shaft providing a shaft axis; anda bearing assembly mounted on each of the gear shafts and providing a bearing axis, the bearing assembly including a spherical bearing configured to permit angular movement of the bearing axis relative to the shaft axis.2. The system according to claim 1 , comprising a gear supported on each bearing assembly for rotation about the bearing axis claim 1 , an input gear located radially inward from and intermeshing with intermediate gears supported on the gear shafts claim 1 , and a ring gear arranged about and intermeshing with the intermediate gears claim 1 , the input gear supported by an input shaft claim 1 , and fixed structure supporting one of the ring gear and the torque frame claim 1 , and the other of the ring gear and the torque frame coupled to a fan shaft.3. The system according to claim 2 , wherein the system is a star gear system in which the ring gear is coupled to the fan shaft and the torque frame is secured to the fixed structure.4. The system according to claim 1 , wherein the bearing assembly includes a race supporting the gear claim 1 , and the spherical bearing is received by the race claim 1 , and a pin configured to prevent relative rotation between the race claim 1 , spherical bearing and the shaft about the shaft axis.5. The system according to claim 1 , wherein the torque frame includes an oil passage provided through the shaft and configured to provide lubricating ...

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Номер записи: 71
03-01-2013 дата публикации

Dual fan gas turbine engine and gear train

Номер: US20130004297A1
Автор: William G. Sheridan
Принадлежит: Individual

A gas turbine engine includes a first fan, a second fan spaced axially from the first fan, a turbine-driven fan shaft and an epicyclic gear train coupled to be driven by the turbine-driven fan shaft and coupled to drive the first fan and the second fan. The epicyclic gear train includes a carrier that supports star gears that mesh with a sun gear, and a ring gear that surrounds and meshes with the star gears. The star gears are supported on respective journal bearings.

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Номер записи: 72
17-01-2013 дата публикации

DUAL MODEL SCAVENGE SCOOP

Номер: US20130016936A1
Автор: Glahn Jorn A., James Denman H., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A system for removing oil from a bearing compartment has a port connected to an end wall of the compartment through which the oil exits the compartment, a scavenge scoop connected to the port for collecting the oil, and a separation device connected to the scavenge scoop for creating an oil collection region. 110-. (canceled)11. A bearing compartment comprising:a bearing;means for introducing an airflow into said compartment;means for introducing a flow of oil into said compartment to lubricate said bearing and cool said compartment;means for introducing an airflow into said compartment to reduce the leakage of any oil from the compartment;means for removing said oil from said compartment, said oil removing means comprising a port connected to an end wall of said compartment through which said oil exits said compartment, a scavenge scoop connected to said port for collecting oil, and a separation device connected to said scavenge scoop for creating an oil collection region.12. The bearing compartment according to claim 11 , wherein said scavenge scoop is a tangential scavenge scoop.13. The bearing compartment according to claim 11 , wherein said scavenge scoop has a first wall and a second wall at an angle to said first wall.14. The bearing compartment according to claim 13 , further comprising said port having an exit pipe and said first wall extending into said exit pipe.15. The bearing compartment according to claim 13 , wherein said exit pipe receives oil from both sides of said scoop.16. The bearing compartment according to claim 13 , further comprising said port having an exit pipe claim 13 , said first wall terminating at a distance from an entrance to said exit pipe claim 13 , and a baffle mounted to said end wall.17. The bearing compartment according to claim 16 , wherein said exit pipe receives oil from both sides of said scoop.18. The bearing compartment according to claim 11 , wherein said separation device comprises a separation wall for forming a sump ...

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Номер записи: 73
24-01-2013 дата публикации

Method of Assembly for Gas Turbine Fan Drive Gear System

Номер: US20130023378A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A method of assembling an epicyclic gear train includes the steps of providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced apart apertures provided at an outer circumference of the carrier. Gear pockets are provided between the walls and extend to the apertures, and a central opening in at least one of the walls. A plurality of intermediate gears is inserted through the central opening and move the intermediate gears radially outwardly into the gear pockets to extend through the apertures. A sun gear is inserted through the central opening. The plurality of intermediate gears is moved radially inwardly to engage the sun gear. A gear reduction is also disclosed. 1. A method of assembling an epicyclic gear train comprising the steps of:a) providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining spaced apart apertures provided at an outer circumference of the carrier, gear pockets provided between the walls and extending to the apertures, and a central opening in at least one of the walls;b) inserting a plurality of intermediate gears through the central opening and moving the intermediate gears radially outwardly into the gear pockets to extend into the apertures;c) inserting a sun gear through the central opening; andd) moving the plurality of intermediate gears radially inwardly to engage the sun gear.2. The method as set forth in claim 1 , wherein step d) occurs after step c).3. The method as set forth in claim 1 , wherein journal bearings are inserted within each of said intermediate gears after step d).4. The method as set forth in claim 1 , wherein a ring gear is subsequently placed on an outer periphery of the sun gears to engage the sun gears.5. The method as set forth in claim 4 , wherein said sun gear and said intermediate gears are each formed as a single gear claim 4 , and said ring gear is formed as a two-part gear ...

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Номер записи: 74
28-02-2013 дата публикации

TORQUE FRAME AND ASYMMETRIC JOURNAL BEARING FOR FAN DRIVE GEAR SYSTEM

Номер: US20130051984A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A fan drive gear system for a gas turbine engine includes a torque frame and a journal bearing. The torque frame has an arm that extends therefrom, the journal bearing is mounted on the arm. When the fan drive gear system is in an unloaded condition an outer circumferential surface of the journal bearing is disposed at an angle so that it is non-parallel with respect to an axis of the arm. When the fan drive gear system is in a loaded condition the outer circumferential surface is disposed substantially parallel to the axis of the arm. In another aspect, the journal bearing includes a main body, a first wing, and a second wing. The first and second wings extend from the main body and are asymmetric with respect to one another. 1. A fan drive gear system for a gas turbine engine , comprising:a torque frame having an arm extending therefrom, the arm providing an axis; anda journal bearing mounted on the arm and having an outer circumferential surface;wherein when the fan drive gear system is in an unloaded condition the outer circumferential surface is disposed at an angle that is non-parallel with respect to the axis of the arm, and wherein when the fan drive gear system is in a loaded condition the outer circumferential surface is disposed substantially parallel with respect to the axis of the arm.2. The fan drive gear system of claim 1 , wherein the arm has an outer surface that forms a bias angle with respect to the axis of the arm.3. The fan drive gear system of claim 2 , wherein the angle of the outer surface of the journal bearing is substantially parallel to the bias angle of the outer surface of the arm.4. The fan drive gear system of claim 1 , wherein the journal bearing comprises:a main body having an inner radial surface that interfaces with the arm;a first wing extending from a first side of the main body; anda second wing extending from a second opposing side of the main body, the first and second wings being asymmetric with respect to one another.5. The ...

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Номер записи: 75
25-04-2013 дата публикации

Split accessory drive system

Номер: US20130098058A1
Автор: William G. Sheridan
Принадлежит: United Technologies Corp

A gas turbine engine includes a spool, a first accessory gearbox, a second accessory gearbox, and a scavenge pump. The first accessory gearbox is connected to and driven by the spool. The second accessory gearbox is connected to and driven by the first accessory gearbox. The scavenge pump is connected between the first accessory gearbox and the second accessory gearbox. The first accessory gearbox drives the second accessory gearbox through the scavenge pump.

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Номер записи: 76
11-07-2013 дата публикации

Coupling system for a star gear train in a gas turbine engine

Номер: US20130178327A1
Автор: Alessio Pescosolido, Michael E. McCune, William G. Sheridan
Принадлежит: United Technologies Corp

A star gear train for use in a gas turbine engine includes a sun gear, a ring gear, a plurality of star gears and a coupling system. The sun gear is rotatable by a shaft. The ring gear is secured to a ring gear shaft. Each of the plurality of star gears is rotatably mounted in a star carrier and meshes with the sun gear and the ring gear. The coupling system comprises a sun gear flexible coupling, a carrier flexible coupling and a deflection limiter. The sun gear flexible coupling connects the sun gear to the shaft. The carrier flexible coupling connects the carrier to a non-rotating mechanical ground. The deflection limiter is connected to the star carrier to limit excessive radial and circumferential displacement of the star gear train.

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Номер записи: 77
25-07-2013 дата публикации

GEARED TURBOMACHINE FAN AND COMPRESSOR ROTATION

Номер: US20130186058A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A high-bypass ratio geared turbomachine comprises a compressor section of a high-bypass ratio geared turbomachine. The compressor section provides at least a low-pressure compressor and a high-pressure compressor, wherein a rotor of the low-pressure compressor rotates together with a rotor of a fan. 1. A high-bypass ratio geared turbomachine , comprising:a compressor section of a high-bypass ratio geared turbomachine, the compressor section providing at least a low-pressure compressor and a high-pressure compressor, wherein a rotor of the low-pressure compressor rotates together with a rotor of a fan.2. The high-bypass ratio geared turbomachine of claim 1 , wherein the rotor of the low pressure compressor and the rotor of the fan rotate at the same speed and in the same direction.3. The high-bypass ratio geared turbomachine of claim 1 , wherein the high-bypass ratio geared turbomachine has a fan bypass ratio greater than about 8.4. The high-bypass ratio geared turbomachine of claim 1 , wherein the high-bypass ratio geared turbomachine has an overall compression ratio greater than about 40.5. The high-bypass ratio geared turbomachine of claim 1 , wherein the high-pressure compressor has a pressure ratio greater than about 20.6. The high-bypass ratio geared turbomachine of claim 1 , wherein the fan includes a shaft that is rotatably supported by a plurality of tapered bearings.7. The high-bypass ratio geared turbomachine of claim 1 , including a turbine shaft that rotates a geared architecture to rotate the rotor of the low-pressure compressor and the rotor of the fan.8. The high-bypass ratio geared turbomachine of claim 7 , wherein at least one thrust bearing rotatably supports the turbine shaft claim 7 , and the at least one thrust bearing is located axially between the geared architecture and a turbine secured to the turbine shaft.9. The high-bypass ratio geared turbomachine of claim 8 , wherein the at least one thrust bearing is a bi-directional tapered bearing.10 ...

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Номер записи: 78
01-08-2013 дата публикации

TURBOMACHINE FAN CLUTCH

Номер: US20130195603A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

An exemplary turbomachine clutch assembly includes a clutch that moves from a first position to a second position in response to rotation of a turbomachine fan at a speed greater than a threshold speed. The clutch permits rotation of the turbomachine fan in a first direction whether the clutch is in the first position or the second position. The clutch limits rotation of the turbomachine fan in an opposite, second direction when the clutch is in the first position. 1. A turbomachine clutch assembly , comprising:a clutch that moves from a first position to a second position in response to rotation of a turbomachine fan at a speed greater than a threshold speed, whereinthe clutch permits rotation of the turbomachine fan in a first direction whether the clutch is in the first position or the second position, andthe clutch limits rotation of the turbomachine fan in an opposite, second direction when the clutch is in the first position.2. The turbomachine clutch assembly of claim 1 , wherein the turbomachine fan is a ducted fan.3. The turbomachine clutch assembly of claim 1 , wherein the threshold speed is less than an idling speed.4. The turbomachine clutch assembly of claim 1 , wherein the threshold speed is a threshold rotational speed of the fan.5. The turbomachine clutch assembly of claim 1 , wherein the clutch is an entirely mechanical clutch.6. The turbomachine clutch assembly of claim 1 , wherein the clutch moves from the second position to the first position in response to rotation of a turbomachine fan no longer exceeding the threshold speed.7. The turbomachine clutch assembly of claim 1 , including a lubrication system that lubricates the turbomachine fan claim 1 , the lubrication system powered by the turbomachine fan rotating in the first direction.8. The turbomachine clutch assembly of claim 1 , wherein the clutch permits rotation of the turbomachine fan in the second direction when the clutch is in the second position.9. The turbomachine clutch assembly of ...

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Номер записи: 79
01-08-2013 дата публикации

GAS TURBINE ENGINE BEARING ARRANGEMENT INCLUDING AFT BEARING HUB GEOMETRY

Номер: US20130195647A1
Автор: Muldoon Marc J., Reinhardt Gregory E., Savela Gregory M., Sheridan William G.
Принадлежит:

An example gas turbine engine includes a housing and first and second spools coaxial with one another and arranged within the housing. The first spool is arranged within the second spool and extends between forward and aft ends. The aft end extends axially beyond the second spool. First and second bearings support the aft end of the first spool relative to the housing. An example method of assembling the gas turbine engine includes arranging a first spool within a second spool. First and second bearings are mounted between an aft end of the first spool and a fixed housing. 1. A gas turbine engine comprising:a housing;first and second spools coaxial with one another and arranged within the housing, the first spool arranged within the second spool and extending between forward and aft ends, the aft end extending axially beyond the second spool;first and second bearings supporting the aft end of the first spool relative to the housing;a combustor section and high and low pressure turbine sections downstream from the combustor section, the high pressure turbine section mounted on the second spool, and the low pressure turbine section mounted on the first spool, wherein the housing includes a first housing portion arranged axially between the high and low pressure turbine sections, and a second housing portion arranged aft of the low pressure turbine section;the first and second bearings are supported by the second housing portion; andthe second housing portion includes a hub providing an integrated, unitary structure removably secured to an annular flange, the first and second bearings mounted to the hub,wherein the hub is a single piece.2. The gas turbine engine according to claim 1 , wherein the housing is rotationally fixed claim 1 , the first and second bearings each having first and second radial sides claim 1 , the first radial side supported by the housing claim 1 , and the second radial side supporting the aft end.38.-. (canceled)9. The gas turbine engine ...

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Номер записи: 80
29-08-2013 дата публикации

OIL BAFFLE FOR GAS TURBINE FAN DRIVE GEAR SYSTEM

Номер: US20130223994A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbine engine includes a housing supporting compressor, turbine sections, and an epicyclic gear train including a carrier. Sun gear and intermediate gears are arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubrication on the at least one of the sun gear and intermediate gears. 1. A turbine engine comprising:a housing supporting compressor and turbine sections; andan epicyclic gear train including:a carrier;a sun gear and intermediate gears arranged about and intermeshing with the sun gear, the intermediate gears supported by the carrier; anda baffle supported relative to the carrier including a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubrication on the at least one of the sun gear and intermediate gears.2. The turbine engine according to claim 1 , comprising a ring gear intermeshing with the intermediate gears and an output shaft interconnected to the ring gear claim 1 , and an input shaft interconnected to the sun gear.3. The turbine engine according to claim 2 , wherein the carrier is fixed relative to a housing claim 2 , the output shaft drives a turbo fan claim 2 , and the input shaft supports a compressor hub having compressor blades.4. The turbine engine according to claim 1 , wherein the carrier includes spaced apart walls with circumferentially spaced mounts interconnecting the walls claim 1 , the mounts providing circumferentially spaced apart apertures between the mounts at an outer circumference of the carrier claim 1 , the intermediate gears extending through the apertures to intermesh with a ring gear.5. The turbine engine according to claim 4 , comprising a housing supporting a torque frame that is secured to the mounts.6. The turbine engine according to claim 4 , wherein the lubrication passage includes a primary passage ...

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Номер записи: 81
05-12-2013 дата публикации

AUXILIARY OIL SYSTEM FOR NEGATIVE GRAVITY EVENT

Номер: US20130319798A1
Автор: Dolman Paul H., Sheridan William G.
Принадлежит:

A lubrication system includes a bearing compartment. A main reservoir is fluidly connected to the bearing compartment by a main supply passage. A main pump is arranged in the main supply passage configured to provide fluid from the main reservoir to the bearing compartment during a positive gravity condition. A secondary supply passage fluidly connects the main reservoir to at least one segment of the main supply passage, thereby providing fluid from the main reservoir to the bearing compartment during a negative gravity condition. A method of supplying a bearing compartment with fluid includes pumping a fluid from a main reservoir to a bearing compartment through a main supply passage during a positive gravity condition, and providing fluid from the main reservoir to the bearing compartment through a secondary supply passage, fluidly connected to at least one segment of the main supply passage, in response to a negative gravity condition. 1. A lubrication system for a gas turbine engine comprising:a bearing compartment;a main reservoir fluidly connected to the bearing compartment by a main supply passage, which includes one or more passage segments therein;a main pump arranged in the main supply passage configured to provide fluid from the main reservoir to the bearing compartment during a positive gravity condition; anda secondary supply passage fluidly connecting the main reservoir to at least one segment of the main supply passage, thereby providing fluid from the main reservoir to the bearing compartment during a negative gravity condition.2. The system according to claim 1 , wherein the main reservoir includes upper and lower portions claim 1 , the main supply passage in fluid communication with the lower portion claim 1 , and the secondary supply passage in fluid communication with the upper portion.3. The system according to claim 2 , comprising a first check valve arranged in the secondary supply passage fluidly between the main reservoir and the main ...

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Номер записи: 82
09-01-2014 дата публикации

COUPLING SYSTEM FOR A GEAR TRAIN IN A GAS TURBINE ENGINE

Номер: US20140011623A1
Автор: McCune Michael E., Pescosolido Alessio, Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An engine comprises a gear system and a coupler. The gear system comprises a sun gear, a ring gear, and a plurality of intermediate gears that engage the sun and ring gears. The coupler is for connecting the gear system to an engine static structure. The coupler is configured to limit deflection of the gear system. Two of the sun gear, the ring gear, and the plurality of intermediate gears are configured to rotate and/or orbit about a central axis. The third of the sun gear, the ring gear, and the plurality of intermediate gears is connected to the coupler and is configured not to orbit and/or rotate relative to the central axis. 1. An engine comprising: a sun gear;', 'a ring gear; and', 'a plurality of intermediate gears that engage the sun and ring gears; and, 'a gear system comprisinga coupler for connecting the gear system to an engine static structure, the coupler configured to limit displacement of the gear system;wherein two of the sun gear, the ring gear, and the plurality of intermediate gears are configured to rotate and/or orbit about a central axis; andwherein the third of the sun gear, the ring gear, and the plurality of intermediate gears is connected to the coupler and is configured not to orbit and/or rotate relative to the central axis.2. The engine of wherein the coupler includes:a first member; anda second member that is configured to engage the first member to limit radial deflection of the gear system.3. The engine of claim 2 , wherein the first member is a shoulder and the second member is a flange.4. The engine of wherein the shoulder is configured to engage the flange at a spline to limit circumferential rotation of the gear system relative to the engine static structure.5. The engine of wherein the spline comprises:a plurality of slots extending into the flange; anda plurality of tabs extending from the shoulder to engage the slots.6. The engine of wherein the coupler comprises:a torque frame configured to engage the gear system; anda ...

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Номер записи: 83
23-01-2014 дата публикации

FUNDAMENTAL GEAR SYSTEM ARCHITECTURE

Номер: US20140020404A1
Автор: Ackermann William K., Husband Jason, Kupratis Daniel Bernard, McCune Michael E., Schwarz Frederick M., Sheridan William G., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing energy equal to less than about 2% of energy input into the gear system. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an axis;a compressor section;a combustor in fluid communication with the compressor section;a fan drive turbine in communication with the combustor, wherein the fan drive turbine has a first exit area at a first exit point and is configured to rotate at a first speed;a second turbine section including a second exit area at a second exit point and being configured to rotate at a second speed that is faster than the first speed, wherein a first performance quantity is defined as a product of the first speed squared and the first area, a second performance quantity is defined as a product of the second speed squared and the second area; and a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5;a gear system configured to provide a speed reduction between the fan drive turbine and the fan, and to transfer shaft power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%;a mount flexibly supporting portions of the gear system, the mount extending from a static structure of the engine to accommodate at least radial movement between the gear system and the static structure; anda lubrication system configured to provide lubricant to the gear system and to remove thermal energy from the gear system.2. The gas turbine engine as recited in claim 1 , wherein the lubrication system includes a capacity for ...

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Номер записи: 84
13-02-2014 дата публикации

Method of Assembly for Gas Turbine Fan Drive Gear System

Номер: US20140045645A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A method of mounting a gear train to a torque frame includes providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining mounts for interconnecting the walls. Spaced apart apertures are provided between the mounts at an outer circumference of the carrier. Gear pockets are provided between the walls and mounts extending to the apertures, and a central opening in at least one of the walls. A plurality of intermediate gears and a sun gear are inserted in the carrier. A first ring gear half is placed about the outer periphery of the intermediate gears, and attach a torque frame to the carrier. 1. A method of mounting a gear train to a torque frame comprising:a) providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining mounts for interconnecting the walls, spaced apart apertures provided between the mounts at an outer circumference of the carrier, gear pockets provided between the walls and mounts extending to the apertures, and a central opening in at least one of the walls;b) inserting a plurality of intermediate gears and a sun gear in the carrier; andc) placing a first ring gear half about the outer periphery of said intermediate gears, and attaching a torque frame to said carrier.2. The method as set forth in claim 1 , wherein a second ring gear half is then mounted to the outer periphery subsequent to the torque frame being mounted to the carrier.3. The method as set forth in claim 2 , wherein said torque frame has a plurality of axially extending fingers which are received within slots in the carrier claim 2 , at locations circumferentially intermediate locations of said intermediate gears claim 2 , and said first ring gear half is moved such that it does not block radially inwardly extending apertures in a radially outer surface of said carrier claim 2 , and pins are then moved into said apertures to ...

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Номер записи: 85
06-03-2014 дата публикации

FACE SEAL RETAINING ASSEMBLY FOR GAS TURBINE ENGINE

Номер: US20140062026A1
Автор: Davis Todd A., Sheridan William G.
Принадлежит:

A face seal assembly for a gas turbine engine includes an engine static structure. A guide assembly supports a face seal for movement relative to the engine static structure in an axial direction. The guide assembly includes a guide pin having a first end that supports a washer that is retained by a circlip secured to the first end. The circlip is configured to limit movement of the face seal in the axial direction. 1. A face seal assembly for a gas turbine engine comprising:an engine static structure; anda guide assembly supporting a face seal for movement relative to the engine static structure in an axial direction, the guide assembly including a guide pin having a first end supporting a washer that is retained by a circlip secured to the first end, the circlip configured to limit movement of the face seal in the axial direction.2. The face seal assembly according to claim 1 , wherein the guide pin includes a second end secured to the static structure claim 1 , the axial direction is defined between the first and second ends.3. The face seal assembly according to claim 1 , wherein the face seal includes a carrier slideable relative to the guide pins.4. The face seal assembly according to claim 1 , wherein the face seal includes an annular metal backed carbon seal.5. The face seal assembly according to claim 1 , comprising a rotating structure supported relative to the engine static structure by a bearing claim 1 , and a seal seat mounted on the rotating structure.6. The face seal assembly according to claim 1 , wherein the guide assembly includes a guide provided by a carrier claim 1 , with the guide pin received in the guide.7. The face seal assembly according to claim 1 , wherein the guide assembly includes a sleeve mounted on the guide pin and in sliding engagement with the guide.8. The face seal assembly according to claim 7 , wherein the washer abuts the sleeve.9. The face seal assembly according to claim 8 , wherein the guide pin includes a shoulder claim 8 ...

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Номер записи: 86
27-03-2014 дата публикации

Method for Setting a Gear Ratio of a Fan Drive Gear System of a Gas Turbine Engine

Номер: US20140083107A1
Автор: Hasel Karl L., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan rotatable about an axis and a speed reduction device in communication with the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A fan blade tip speed of the fan is less than 1400 fps. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an axis; anda speed reduction device in communication with the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5,wherein a fan blade tip speed of the fan is less than 1400 fps.2. The gas turbine engine of claim 1 , wherein the speed reduction device includes a star gear system gear ratio of at least 2.6.3. The gas turbine engine of claim 2 , wherein the speed reduction device includes a system gear ratio less than or equal to 4.1.4. The gas turbine engine of claim 3 , including a bypass ratio that is greater than about 6.0.5. A gas turbine engine comprising:a fan section including a fan rotatable about an axis;a speed reduction device in communication with the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 2.6 and less than or equal to 4.1; anda bypass ratio that is between about 11.0 and about 22.0,wherein the speed reduction device includes a star gear system gear ratio of at least 2.6 and less than or equal to 4.1, a bypass ratio that is greater than about 6.0, and a fan blade tip speed of the fan is less than 1400 fps.6. The gas turbine engine of claim 1 , wherein the star system includes a sun gear claim 1 , a plurality of star gears claim 1 , a ring gear claim 1 , and a carrier.7. The gas turbine engine of claim 6 , wherein each of the plurality of star gears include at least one bearing.8. The gas turbine engine of claim 6 , wherein the carrier is fixed from rotation.9. The gas ...

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Номер записи: 87
27-03-2014 дата публикации

GEAR CARRIER FLEX MOUNT LUBRICATION

Номер: US20140087907A1
Автор: Coffin James B., McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An example epicyclic gear train assembly includes a flexure pin received by a carrier. The flexure pin and the carrier respectively include first and second pin apertures configured to receive a retainer pin. The flexure pin further includes a lubricant conduit separate from the first pin aperture. 1. An epicyclic gear train assembly comprising:a flexure pin received by a carrier, the flexure pin and the carrier respectively including at least portions of first and second pin apertures configured to receive a retainer pin, wherein the flexure pin further includes a lubricant conduit separate from the first pin aperture.2. The epicyclic gear train assembly of claim 1 , wherein the lubricant conduit communicates with a bearing lubricant supply that is outside the carrier.3. The epicyclic gear train assembly of claim 2 , wherein the lubricant conduit is a first lubricant conduit claim 2 , and the carrier includes a second lubricant conduit that communicates lubricant between the bearing lubricant supply and the first lubricant conduit.4. The epicyclic gear train assembly of claim 3 , including a metering device that meters flow of lubricant from the second lubricant conduit to the first lubricant conduit.5. The epicyclic gear train assembly of claim 4 , wherein the metering device is partially received within the flexure pin.6. The epicyclic gear train assembly of claim 1 , wherein the lubricant conduit delivers lubricant to an interface between the flexure pin and a bushing.7. The epicyclic gear train assembly of claim 6 , including a torque frame that receives the bushing.8. The epicyclic gear train assembly of claim 1 , wherein the retainer pin is a bolt.9. The epicyclic gear train assembly of claim 1 , wherein the lubricant conduit comprises a first portion extending transverse to an axis of rotation of a gear claim 1 , and a second portion extending parallel to the axis of rotation of the gear.10. The epicyclic gear train assembly of claim 1 , including a clocking ...

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Номер записи: 88
10-04-2014 дата публикации

OIL BAFFLE FOR GAS TURBINE ENGINE FAN DRIVE GEAR SYSTEM

Номер: US20140099187A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

An exemplary turbine engine assembly includes a first shaft that is rotatably driven by a second shaft of a gas turbine engine, a compressor hub driven by the second shaft, the compressor hub within a compressor section of the gas turbine engine, an epicyclic gear train that is driven by the first shaft, and a common attachment point that secures the first shaft and the compressor hub to the second shaft. 1. A turbine engine assembly comprising:a first shaft that is rotatably driven by a second shaft of a gas turbine engine;a compressor hub driven by the second shaft, the compressor hub within a compressor section of the gas turbine engine;an epicyclic gear train that is driven by the first shaft; anda common attachment point that secures the first shaft and the compressor hub to the second shaft.2. The turbine engine assembly of claim 1 , the epicyclic gear train having:a carrier, anda sun gear and intermediate gears arranged about and intermeshing with the sun gear, the intermediate gears supported by the carrier.3. The turbine engine assembly of claim 2 , including a baffle secured to the carrier by a fastening member claim 2 , the baffle including a lubrication passage near at least one of the sun gear and intermediate gears for directing lubricant on the at least one of the sun gear and intermediate gears.4. The turbine engine assembly of claim 2 , comprising a ring gear intermeshing with the intermediate gears and a third shaft interconnected to the ring gear claim 2 , and the first shaft interconnected to the sun gear.5. The turbine engine assembly of claim 4 , wherein the carrier is fixed relative to a housing claim 4 , the third shaft drives a fan claim 4 , and the first shaft supports a compressor hub having compressor blades.6. The turbine engine assembly of claim 1 , wherein the first shaft is a compressor shaft.7. The turbine engine assembly of claim 1 , wherein the epicyclic gear train rotatably drives a third shaft.8. The turbine engine assembly of ...

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Номер записи: 89
06-01-2022 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20220003172A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). A non-rotatable flexible coupling and a rotatable flexible coupling support an epicyclic gear system. The couplings are subject to a Motion II of cantilever beam free end motion with respect to the engine central axis. The non-rotatable and the rotatable flexible couplings each have a stiffness of a common stiffness type under a common type of motion. The common stiffness type is a Stiffness B and the common type of motion is the Motion II. The Stiffness B of the rotatable flexible coupling is greater than the stiffness B of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a lateral frame stiffness (LFS);an epicyclic gear system coupled to the fan shaft; anda non-rotatable flexible coupling and a rotatable flexible coupling supporting the epicyclic gear system, the non-rotatable flexible coupling and the rotatable flexible coupling being subject to a Motion II of cantilever beam free end motion with respect to the engine central axis,the non-rotatable flexible coupling and the rotatable flexible coupling each having a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis, the common stiffness being defined with respect to the LFS, the common stiffness type is a Stiffness B and the common type of motion is the Motion II, the Stiffness B of the rotatable flexible coupling being greater than the stiffness of the non-rotatable flexible coupling, and a ratio of LFS/Stiffness B of the non-rotatable flexible coupling is in a range of 10-40.2. The gas turbine engine as recited ...

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Номер записи: 90
07-01-2016 дата публикации

GEARED TURBOFAN WITH GEARBOX SEAL

Номер: US20160003142A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan, a compressor section, a turbine section, and a gear reduction for driving the fan through the turbine section. A rotating element and at least one bearing compartment includes a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment, and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. The seal has a plurality of sealing members extending radially toward a sealing surface. 1. A gas turbine engine comprising:a fan, a compressor section, a turbine section, and a gear reduction for driving said fan through said turbine section;a rotating element and at least one bearing compartment including a bearing for supporting said rotating element, a seal for resisting leakage of lubricant outwardly of said bearing compartment, and for allowing pressurized air to flow from a chamber adjacent said seal into the bearing compartment; andsaid seal having a plurality of sealing members extending radially toward a sealing surface.2. The gas turbine engine as set forth in claim 1 , wherein said seal is a labyrinth seal having a plurality of knife edges.3. The gas turbine engine as set forth in claim 2 , wherein a first radius is defined to a radial extent of said knife edges and a second radius may be defined on a drive shaft associated with said fan drive turbine at a location in a plane defined by a fuel nozzle in a combustor in said gas turbine engine claim 2 , and a diameter ratio of said first radius to said second radius being less than or equal to about 2.0.4. The gas turbine engine as set forth in claim 3 , wherein said diameter radius being less than or equal to about 1.75.5. The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with said gear reduction.6. The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with said fan.7. The gas ...

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Номер записи: 91
07-01-2016 дата публикации

TURBOMACHINE FAN CLUTCH

Номер: US20160003143A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine assembly includes, among other things, a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed. Whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction. When the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction. The clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure. The clutch is removable on-wing. 1. A gas turbine engine assembly , comprising:a clutch configured to move from a first position to a second position in response to rotation of a gas turbine engine fan at a speed greater than a threshold speed,wherein, whether the clutch is in the first position or the second position, the clutch permits rotation of the gas turbine engine fan in a first direction, and when the clutch is in the first position, the clutch limits rotation of the gas turbine engine fan only in an opposite, second direction,wherein the clutch is disposed within a compartment that is accessible and removable via removal of an aft engine cover structure, whereby the clutch is removable on-wing.2. The gas turbine engine assembly of claim 1 , wherein the aft engine cover structure includes an engine exhaust cone.3. The gas turbine engine assembly of claim 2 , wherein the clutch is disposed within an aft bearing compartment and the aft engine cover structure further includes an aft bearing compartment cover plate claim 2 , disposed axially inward of the exhaust cone.4. The gas turbine engine assembly of claim 1 , wherein the clutch is positioned within a gas turbine engine such that the clutch can be moved from an installed position within the gas turbine engine to an uninstalled position without removing any blades from the gas turbine ...

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Номер записи: 92
27-01-2022 дата публикации

Hybrid electric fan with stall free low pressure compressor

Номер: US20220025823A1
Автор: Christopher J. Hanlon, Daniel Bernard Kupratis, William G. Sheridan
Принадлежит: Raytheon Technologies Corp

A gas turbine engine according to an exemplary embodiment of this disclosure includes among other possible things, a fan section including a plurality of fan blades, a first electric motor assembly that provides a first drive input for driving the fan blades about an axis, a turbine section, and a geared architecture driven by the turbine section and coupled to the fan section to provide a second drive input for driving the fan blades, and second electric motor assembly is coupled to rotate the geared architecture relative to a fixed structure controls a speed of the fan blades provided by a combination of the first drive input and the second drive input.

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Номер записи: 93
14-01-2016 дата публикации

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

Номер: US20160010562A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section including a fan hub. A speed reduction device includes a star gear system. A turbine section is connected to the fan section through the speed reduction device. A first fan bearing for supporting rotation of the fan hub is connected forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is connected aft of the speed reduction device. A first outer race of the first fan bearing is attached to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan hub;a speed reduction device including a star gear system;a turbine section connected to the fan section through the speed reduction device;a first fan bearing for supporting rotation of the fan hub connected forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub connected aft of the speed reduction device; anda first outer race of the first fan bearing attached to the fan hub.2. The gas turbine engine of including a compressor section configured to rotate with the fan section.3. The gas turbine engine of including a first inner race of the first fan bearing connected to a static structure and a second inner race of the second fan bearing connected to a static structure.4. The gas turbine engine of wherein the first bearing and the second bearing include at least one of roller bearings claim 1 , ball bearings claim 1 , or tapered bearings.5. The gas turbine engine of including a high pressure compressor with a compression ratio of at least 20:1.6. The gas turbine engine of including a low pressure compressor with a compression ratio of at least 2:1.7. The gas turbine engine of including a fan by pass ratio greater than 10.8. The gas turbine engine of claim 1 , wherein the star gear system includes a sun gear claim 1 , star gears claim 1 , a ring gear mechanically attached to the fan section claim 1 , and a carrier fixed from rotation.9. The gas turbine engine of claim 8 , ...

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Номер записи: 94
14-01-2016 дата публикации

OIL LOSS PROTECTION FOR A FAN DRIVE GEAR SYSTEM

Номер: US20160010563A1
Автор: Sheridan William G.
Принадлежит:

A fan drive gear system includes at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating. At least a first axial end of the intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system. The fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event. 1. A fan drive gear system comprising:at least one intermediate gear that includes an axial gear passage for receiving and conveying a fluid suitable for cooling and/or lubricating;at least a first axial end of said intermediate gear includes a first fluid storage trap for capturing fluid entering and/or exiting the gear passage and storing the fluid therein during powered operation of the fan drive gear system; andwhereby the fluid is capable of being passively supplied to the intermediate gear passage during an interrupted power event.2. The fan drive gear system of claim 1 , further comprising;a sun gear interfaced with said intermediate gear;a ring gear interfaced with said intermediate gear; anda carrier body supporting the intermediate gear.3. The fan drive gear system of claim 2 , wherein the at least one fluid trap comprises a radially outward base portion relative to an axis defined by said carrier body and at least one radially inward base portion relative to said axis defined by said carrier body.4. The fan drive gear system of claim 3 , wherein said radially outward base portion is defined on a first axial end by a radially aligned wall segment of said carrier body relative to the axis defined by the carrier body and said radially outward base portion is defined on a second axial end by a radially aligned wall of said trap.5. The fan drive gear system of claim 4 , wherein said radially aligned wall of said fluid storage trap extends ...

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Номер записи: 95
11-01-2018 дата публикации

GAS TURBINE ENGINE STARTER REDUCTION GEAR TRAIN WITH GEARED ROTARY ACTUATOR

Номер: US20180010523A1
Автор: Brown Jeff A., Davis Richard Alan, Fox Paul F., Harder Benjamin T., Hergert Richard R., McMasters Brian, Nguyen James Vandung, Roberts Jeffrey Todd, Sheridan William G., Slayter Matthew Allen, Walker Daniel Richard, Wilson Dwayne Leon
Принадлежит:

According to an aspect, a system for a gas turbine engine includes a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine. The reduction gear train includes a starter interface gear that engages the starter gear train and a core-turning clutch operably connected to the starter interface gear. The reduction gear train also includes a geared rotary actuator including a primary planetary gear system, where the geared rotary actuator is operably connected to the core-turning clutch. The reduction gear train further includes a secondary planetary gear system operably connected to the primary planetary gear system and a core-turning input. The system also includes a mounting pad with an interface to couple a core-turning motor to the core-turning input of the reduction gear train. 1. A system for a gas turbine engine comprising: a starter interface gear that engages the starter gear train;', 'a core-turning clutch operably connected to the starter interface gear;', 'a geared rotary actuator comprising a primary planetary gear system, the geared rotary actuator operably connected to the core-turning clutch; and', 'a secondary planetary gear system operably connected to the primary planetary gear system and a core-turning input; and, 'a reduction gear train operable to drive rotation of a starter gear train that interfaces to an accessory gearbox of the gas turbine engine, the reduction gear train comprisinga mounting pad comprising an interface to couple a core-turning motor to the core-turning input of the reduction gear train.2. The system of claim 1 , wherein the starter interface gear engages a planet gear of the starter gear train claim 1 , and the starter gear train is operably connected to the accessory gearbox through a starter clutch.3. The system of claim 1 , wherein the core-turning clutch is an overrunning clutch.4. The system of claim 1 , wherein the geared rotary actuator ...

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Номер записи: 96
11-01-2018 дата публикации

GEARED GAS TURBINE ENGINE

Номер: US20180010551A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section that includes a fan rotatable about an engine axis. A compressor section includes a low pressure compressor rotatable about the engine axis. A turbine section includes a fan drive turbine for driving the fan and the low pressure compressor. A speed reduction device connects the fan drive turbine to the fan and the low pressure compressor. The speed reduction device includes a sun gear driven by an inner shaft. A plurality of intermediate gears surround the sun gear. A carrier supports the plurality of intermediate gears for driving the low pressure compressor. A ring gear is located radially outward from the intermediate gears and includes a forward portion for driving a fan drive shaft and an aft portion. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an engine axis;a compressor section including a low pressure compressor rotatable about the engine axis;a turbine section including a fan drive turbine for driving the fan and the low pressure compressor; and a sun gear driven by an inner shaft;', 'a plurality of intermediate gears surrounding the sun gear;', 'a carrier supporting the plurality of intermediate gears for driving the low pressure compressor; and', 'a ring gear located radially outward from the intermediate gears including a forward portion for driving a fan drive shaft and an aft portion., 'a speed reduction device connecting the fan drive turbine to the fan and the low pressure compressor, the speed reduction device including2. The gas turbine engine of claim 1 , wherein the carrier includes an axially forward portion and an axially aft portion claim 1 , the plurality of intermediate gears are located axially between the axially forward portion and the axially aft portion of the carrier.3. The gas turbine engine of claim 2 , wherein axially aft portion of the carrier is connected to the low pressure compressor.4. The gas turbine engine of claim 1 , wherein the fan drive ...

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Номер записи: 97
14-01-2021 дата публикации

GEARED TURBOFAN WITH FOUR STAR/PLANETARY GEAR REDUCTION

Номер: US20210010386A1
Автор: McCune Michael E., Schwarz Frederick M., Sheridan William G., Suciu Gabriel L.
Принадлежит:

A turbofan engine assembly includes a nacelle and a turbofan engine. The turbofan engine includes a fan, which includes a fan rotor having fan blades, and a nacelle enclosing the fan rotor and the blades. A turbine rotor drives the fan rotor. An epicyclic gear reduction is positioned between the fan rotor and the turbine rotor. The epicyclic gear reduction includes a ring gear, a sun gear, and four intermediate gears that engage the sun gear and the ring gear. A gear ratio of the gear reduction is greater than 3.06. The fan drive turbine drives the sun gear to, in turn, drive the fan rotor. 1. A turbofan engine comprising:a fan including a fan rotor having fan blades, and housing enclosing said fan rotor and said blades;a turbine rotor driving said fan rotor;an epicyclic gear reduction positioned between said fan rotor and said turbine rotor, said epicyclic gear reduction including a ring gear, a sun gear, and four intermediate gears that engage said sun gear and said ring gear, a gear ratio of said gear reduction is greater than 3.06;wherein said fan drive turbine drives said sun gear to, in turn, drive said fan rotor;a bypass ratio is defined as a volume of air delivered by said fan into a bypass duct inward of said housing compared to a volume of air delivered into a compressor, said bypass ratio is greater than or equal to 12.0; andwherein there is a primary oil supply supplying oil to journal bearings that support said intermediate gears.2. The turbofan engine as set forth in claim 1 , wherein said gear ratio is greater than or equal to 4.0.3. The turbofan engine as set forth in claim 2 , wherein said gear ratio is greater than or equal to 4.2.4. The turbofan engine as set forth in claim 2 , wherein said gear ratio is less than or equal to 4.4.5. The turbofan engine as set forth in claim 2 , wherein there is an auxiliary oil supply supplying oil to said journal bearings when there is windmilling of said fan rotor.6. The turbofan engine as set forth in claim 1 , ...

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Номер записи: 98
14-01-2021 дата публикации

GEAR REDUCTION FOR LOWER THRUST GEARED TURBOFAN

Номер: US20210010426A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan rotor having a hub and a plurality of fan blades extending radially outwardly of the hub. A compressor is positioned downstream of the fan rotor, and has a first compressor blade row defined along a rotational axis of the fan rotor and the compressor rotor. A gear reduction is positioned axially between the first compressor blade row and the fan rotor, and includes a ring gear and a carrier. The carrier has an axial length and the ring gear has an outer diameter. A ratio of the axial length to the outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40. The gear reduction is connected to drive the hub to rotate. A method of designing a gas turbine engine is also disclosed. 1. A gas turbine engine comprising:a fan rotor having a hub and a plurality of fan blades extending radially outwardly of said hub,a compressor positioned downstream of the fan rotor, the compressor having a first compressor blade row defined along a rotational axis of said fan rotor and said compressor rotor, anda gear reduction positioned axially between said first compressor blade row and said fan rotor, said gear reduction including a ring gear and a carrier, said carrier having an axial length and said ring gear having an outer diameter, wherein a ratio of said axial length to said outer diameter may be greater than or equal to about 0.20 and less than or equal to about 0.40, and wherein said gear reduction is connected to drive said hub to rotate.2. The gas turbine engine as set forth in claim 1 , wherein a volume is defined for said carrier and said ring gear claim 1 , and said volume being greater than or equal to about 899 inchesand less than or equal to about 1349 inches.3. The gas turbine engine as set forth in claim 1 , wherein the hub has a radius defined at an inlet point of said hub claim 1 , wherein said fan blades have a radius claim 1 , and a ratio of said hub radius to said fan blade radius is less than ...

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Номер записи: 99
14-01-2021 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20210010428A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). An epicyclic gear system is coupled to the fan shaft, and a non-rotatable flexible coupling and a rotatable flexible coupling support the epicyclic gear system. The non-rotatable flexible coupling and the rotatable flexible coupling each have a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis. The stiffness is defined with respect to the LFS. The stiffness of the rotatable flexible coupling is greater than the stiffness of the non-rotatable flexible coupling. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a lateral frame stiffness (LFS);an epicyclic gear system coupled to the fan shaft; anda non-rotatable flexible coupling and a rotatable flexible coupling supporting the epicyclic gear system,the non-rotatable flexible coupling and the rotatable flexible coupling each having a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis, the stiffness being defined with respect to the LFS, the stiffness of the rotatable flexible coupling being greater than the stiffness of the non-rotatable flexible coupling.2. The gas turbine engine as recited in claim 1 , wherein the common type of motion is selected from Motion I claim 1 , Motion II claim 1 , Motion III claim 1 , or Motion IV claim 1 , where Motion I is parallel offset guided end motion claim 1 , Motion II is cantilever beam free end motion and Motion III is angular misalignment no offset motion and Motion IV is axial motion.3. The gas turbine engine as recited in claim 2 , wherein the epicyclic gear system includes a sun gear in meshed engagement with multiple intermediate gears that are ...

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Номер записи: 100
03-02-2022 дата публикации

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

Номер: US20220034263A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine includes a nacelle, and a bypass flow path in a bypass duct within the nacelle of the turbofan engine. A fan section includes a fan with fan blades. The fan section drives air along the bypass flow path. A fan shaft drives a fan that has fan blades and the fan rotates about a central longitudinal axis of the turbofan engine. A speed reduction device includes an epicyclic gear system. A turbine section is connected to the fan section through the speed reduction device and the turbine section rotates about the central longitudinal axis. A first fan bearing for supporting rotation of the fan hub is located axially forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is located axially aft of the speed reduction device. A first outer race of the first fan bearing is fixed relative to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan having fan blades extending from a fan hub, wherein said fan section drives air along a bypass flow path;a fan shaft driving said fan and said fan rotates about a central longitudinal axis of said gas turbine engine;a speed reduction device including an epicyclic gear system including a plurality of intermediate gears supported on a corresponding one of a plurality of flexible posts on a static carrier of the epicyclic gear system;a turbine section connected to the fan section through the speed reduction device and said turbine section rotates about said central longitudinal axis;a first fan bearing for supporting rotation of the fan hub located axially forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub located axially aft of the speed reduction device; anda first outer race of the first fan bearing is fixed relative to the fan hub.2. The gas turbine engine of claim 1 , wherein each of the plurality of flexible posts extend from the static carrier at a proximal end to a distal free end.3. The gas ...

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Номер записи: 101
21-01-2016 дата публикации

GEARED GAS TURBINE ENGINE WITH OIL DEAERATOR

Номер: US20160017812A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction. The lubrication system includes a lubricant pump supplying a mixed air and oil to a deaerator inlet. The deaerator includes a separator that for separating oil, and delivering separated air to an air outlet, and for delivering separated oil back into an oil tank. The separator includes a member having lubricant flow paths on both of two opposed sides. A method of designing a gas turbine engine is also disclosed.

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Номер записи: 102
17-01-2019 дата публикации

GAS TURBINE ENGINE WITH GEARBOX HEALTH FEATURES

Номер: US20190017410A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan and a braking system. The braking system is configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan. Further, when the rotation of the fan sufficiently slows, the braking system is further configured to apply a second level of braking more restrictive than the first level of braking. 1. A gas turbine engine , comprising:a fan; anda lubrication system configured to pump lubricant into a fan drive gearbox when the fan is windmilling at any rotational speed and direction.2. The gas turbine engine as recited in claim 1 , wherein the lubrication system is configured to pump lubricant to the fan drive gearbox when the fan rotates below 1 claim 1 ,000 rpm.3. The gas turbine engine as recited in claim 1 , wherein:the lubrication system includes a main pump and a main reservoir fluidly coupled to the main pump, the main pump configured to pump lubricant from the main reservoir to the fan drive gearbox during normal operating conditions; andthe lubrication system further includes a secondary pump and a secondary reservoir fluidly coupled to the secondary pump, the secondary pump configured to pump lubricant from the secondary reservoir to the fan drive gearbox when the main pump does not provide adequate lubricant to the fan drive gearbox.4. The gas turbine engine as recited in claim 3 , wherein the secondary pump is one of (1) an accessory pump claim 3 , (2) a rotary-shaft driven pump claim 3 , (3) an electrically-driven pump claim 3 , and (4) an aircraft hydraulic system-powered pump.5. A gas turbine engine claim 3 , comprising:a fan;a geared architecture coupled to the fan;at least one sensor configured to generate signals indicative of a condition of the geared architecture; anda control unit electrically coupled to the at least one sensor, the control unit configured to interpret ...

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Номер записи: 103
28-01-2016 дата публикации

Turbofan Engine Bearing and Gearbox Arrangement

Номер: US20160025003A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine () comprising a fan shaft () configured to rotate about an axis () of the engine. A fan drive gear system () is configured to drive the fan shaft. The fan drive gear system has a centerplane (). A first spool comprises a high pressure turbine () and a high pressure compressor (). A second spool comprises an intermediate pressure turbine (), a lower pressure compressor, and a shaft () coupling the intermediate pressure turbine to the lower pressure compressor. A third spool comprises a lower pressure turbine () coupled to the fan drive gear system to drive the fan. The engine has a plurality of main bearings. The turbofan engine has a single stage fan. Of the main bearings, at least one is a shaft-engaging bearing engaging a driving shaft () coupled to the fan drive gear system. A closest () of the shaft-engaging bearings engaging the driving shaft behind the fan drive gear system has a centerplane () and a characteristic radius (RB). The half angle (θ) of a virtual cone () intersecting the core flowpath inboard boundary at the gear system centerplane () and said closest of the shaft-engaging bearings at the characteristic radius (RB) is greater than about 32°. A hub-to-tip ratio (HR:FR) of the fan is less than about 0.38. 2. The engine of wherein:{'sub': 'D', 'b': 540', '580, 'a length Lbetween the centerplane () and a centerplane () of a forward/upstreammost compressor disk is at least one of{'sub': 'G', 'about 2.0 times or less of a gear width Lof the fan drive gear system;'}{'sub': 'T2', 'b': '580', 'about 60% or less of a core flowpath inboard radius Rat the forward/upstreammost compressor disk centerplane (); and'}{'sub': 'T', 'b': '540', 'about 50% less of a core flowpath inboard radius Rat the centerplane ().'}3. The engine of wherein:{'sub': 'D', 'b': 540', '580, 'claim-text': [{'sub': 'G', 'about 1.5 times or less of the gear width Lof the fan drive gear system;'}, {'sub': 'T2', 'b': '580', 'about 50% or less of the core flowpath inboard ...

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Номер записи: 104
23-01-2020 дата публикации

AIRCRAFT HYBRID PROPULSION FAN DRIVE GEAR SYSTEM DC MOTORS AND GENERATORS

Номер: US20200023982A1
Автор: Kupratis Daniel Bernard, Levasseur Glenn, Schwarz Frederick M., Sheridan William G.
Принадлежит:

An aircraft propulsion system is disclosed and includes a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan, a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan and a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors. 1. An aircraft propulsion system comprising:a first gas turbine engine including a first input shaft driving a first gear system, a first fan driven by the first gear system, a first generator supported on the first input shaft and a fan drive electric motor providing a drive input to the first fan;a second gas turbine engine including a second input shaft driving a second gear system, a second fan driven by the second gear system, a second generator supported on the second input shaft and a second fan drive electric motor providing a drive input to the second fan;a controller controlling power output from each of the first and second generators and directing the power output between each of the first and second fan drive electric motors.2. The aircraft propulsion system as recited in claim 1 , wherein the first and second gear systems provide a main drive input to a corresponding one of the first and second fans and the first and second fan drive electric motors provide a supplemental drive input to at least one of the first and second fans.3. The aircraft propulsion system as recited in claim 1 , wherein the controller is configured to balance power generated by both the first and second generators between the ...

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Номер записи: 105
23-01-2020 дата публикации

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

Номер: US20200025101A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine includes a nacelle, and a bypass flow path in a bypass duct within the nacelle of the turbofan engine. A fan section includes a fan with fan blades. The fan section drives air along the bypass flow path. A fan shaft drives a fan that has fan blades and the fan rotates about a central longitudinal axis of the turbofan engine. A speed reduction device includes an epicyclic gear system. A turbine section is connected to the fan section through the speed reduction device and the turbine section rotates about the central longitudinal axis. A first fan bearing for supporting rotation of the fan hub is located axially forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is located axially aft of the speed reduction device. A first outer race of the first fan bearing is fixed relative to the fan hub. 1. A gas turbine engine comprising:a nacelle, and a bypass flow path in a bypass duct within said nacelle of said turbofan engine;a fan section including a fan with fan blades, wherein said fan section drives air along said bypass flow path a fan shaft driving a fan having fan blades and said fan rotates about a central longitudinal axis of said turbofan engine;a speed reduction device including an epicyclic gear system;a turbine section connected to the fan section through the speed reduction device and said turbine section rotates about said central longitudinal axis;a first fan bearing for supporting rotation of the fan hub located axially forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub located axially aft of the speed reduction device; anda first outer race of the first fan bearing fixed relative to the fan hub.2. The gas turbine engine of claim 1 , wherein the second fan bearing includes an outer race and the outer race of the first fan bearing and the outer race of the second fan bearing are fixed relative to the fan hub and rotate with the fan hub in the ...

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Номер записи: 106
04-02-2016 дата публикации

GEARED GAS TURBINE ENGINE WITH OIL DEAERATOR AND AIR REMOVAL

Номер: US20160032770A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine has a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction, and includes a lubricant pump to supply an air/oil mixture to an inlet of a deaerator. The deaerator includes a separator for separating oil and air, delivering separated air to an air outlet, and delivering separated oil back into an oil tank. The separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank. Air within the oil tank moves outwardly through an air exit into the deaerator. A method of designing a gas turbine engine is also disclosed. 1. A gas turbine engine comprising:a fan drive turbine for driving a gear reduction, said gear reduction for driving a fan rotor; and separating oil and air,', 'delivering separated air to an air outlet, and', 'delivering separated oil back into an oil tank,, 'a lubrication system for supplying oil to said gear reduction, the lubrication system including a lubricant pump supplying an air/oil mixture to an inlet of a deaerator, said deaerator including a separator forwherein said separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank, andwherein air within the oil tank moves outwardly through an air exit into the deaerator.2. The gas turbine engine as set forth in claim 1 , wherein the deaerator has an air outlet claim 1 , and an exit guide extending into the deaerator from the air outlet claim 1 , and the deaerator inlet delivering the air/oil mixture about the exit guide claim 1 , and against a wall of the deaerator such that air and oil are separated.3. The gas turbine engine as set forth in claim 1 , wherein the separated oil enters the oil tank through a diffusor.4. The gas turbine engine as set forth in claim 1 , wherein an inlet velocity to the deaerator is less than or equal to 14 feet/second ...

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Номер записи: 107
04-02-2016 дата публикации

Self Cleaning Debris Filter for Fan Drive Gear System

Номер: US20160032772A1
Автор: Goetschius Alan J., Sheridan William G.
Принадлежит:

A pump system for a gas turbine engine has at least one pump. At least one valve has an outlet and at least one inlet is fluidly connected to the at least one pump. A geared architecture is positioned within a bearing compartment. The geared architecture is configured to receive lubricating fluid from the outlet of the at least one valve, a self-cleaning filter is positioned downstream of the at least one valve and upstream of the geared architecture. A gas turbine engine and a method are also disclosed. 1. A pump system for a gas turbine engine comprising:at least one pump;at least one valve having an outlet and at least one inlet fluidly connected to the at least one pump;a geared architecture positioned within a bearing compartment, wherein the geared architecture is configured to receive lubricating fluid from the outlet of the at least one valve; anda self-cleaning filter positioned downstream of the at least one valve and upstream of the geared architecture.2. The pump system according to wherein the self-cleaning filter comprises a filter screen positioned within a filter housing claim 1 , and wherein the filter housing has a filter inlet fluidly connected to the outlet of the at least one valve claim 1 , a first outlet fluidly connected to the geared architecture claim 1 , and a second outlet fluidly connected to the bearing compartment.3. The pump system according to wherein fluid flows into the filter inlet at a first flow rate and flows out of the first outlet at a flow rate that is less than the first flow rate.4. The pump system according to wherein fluid flows out of the second outlet at a third flow rate that is less than the second flow rate.5. The pump system according to wherein the second flow rate is approximately 95% of the first flow rate.6. The pump system according to wherein the filter screen is set to filter to a level comprising approximately 75 microns.7. The pump system according to wherein the at least one inlet of the at least one ...

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Номер записи: 108
04-02-2016 дата публикации

Turbofan Engine Bearing and Gearbox Arrangement

Номер: US20160032827A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine () has a fan shaft () coupling a fan drive gear system () to the fan (). A low spool comprises a low pressure turbine () and a low shaft () coupling the low pressure turbine to the fan drive gear system. A core spool comprises a high pressure turbine (), a compressor (), and a core shaft () coupling the high pressure turbine to the core spool compressor. A first bearing () engages the fan shaft, the first bearing being a thrust bearing. A second bearing () engages the fan shaft on an opposite side of the fan drive gear system from the first bearing, the second bearing being a roller bearing. A third bearing () engages the low spool shaft and the fan shaft. 120. A turbofan engine () comprising:{'b': '28', 'a fan ();'}{'b': '60', 'a fan drive gear system ();'}{'b': '120', 'a fan shaft () coupling the fan drive gear system to the fan;'} [{'b': '50', 'a low pressure turbine (); and'}, {'b': '56', 'a low shaft () coupling the low pressure turbine to the fan drive gear system;'}], 'a low spool comprising [{'b': '46', 'a high pressure turbine ();'}, {'b': '44', 'a compressor (); and'}, {'b': '52', 'a core shaft () coupling the high pressure turbine to the core spool compressor;'}, {'b': '150', 'a first bearing () engaging the fan shaft, the first bearing being a thrust bearing;'}, {'b': '160', 'a second bearing () engaging the fan shaft on an opposite side of the fan drive gear system from the first bearing, the second bearing being a roller bearing; and'}, {'b': '180', 'a third bearing () engaging the low spool shaft and the fan shaft.'}], 'a core spool comprising2. The engine of wherein:{'b': '180', 'the third bearing () is a thrust bearing.'}3. The engine of wherein:{'b': '150', 'the first bearing () is a tapered roller bearing oriented to resist at least forward movement of the fan shaft.'}4. The engine of further comprising:{'b': 190', '212, 'a fourth bearing () engaging the low shaft and a fixed frame (); and'}{'b': 230', '54, 'a fifth bearing () ...

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Номер записи: 109
17-02-2022 дата публикации

DEFLECTION LIMITER FOR A GAS TURBINE ENGINE

Номер: US20220049629A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A gas turbine engine includes a turbine section that includes a fan drive turbine. A geared architecture includes a sun gear in driving engagement with the fan drive turbine. A plurality of planet gears surrounds the sun gear. A ring gear surrounds the plurality of planet gears. A deflection limiter mechanically attaches the ring gear to an engine static structure. The deflection limiter includes a first support fixed to the ring gear that has a first interlocking feature and a second support fixed to the engine static structure that has a second interlocking feature. The first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm). A fan section includes a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft. 1. A gas turbine engine comprising:a turbine section including a fan drive turbine;{'claim-text': ['a sun gear in driving engagement with the fan drive turbine;', 'a plurality of planet gears surrounding the sun gear; and', 'a ring gear surrounding the plurality of planet gears;'], '#text': 'a geared architecture including:'}a deflection limiter mechanically attaching the ring gear to an engine static structure, wherein the deflection limiter includes a first support fixed to the ring gear having a first interlocking feature and a second support fixed to the engine static structure having a second interlocking feature, wherein the first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm); anda fan section including a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft.2. The gas turbine engine of claim 1 , wherein ...

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Номер записи: 110
31-01-2019 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20190032570A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, a gear system rotatably coupled to the fan shaft, and first and second flexible couplings supporting the gear system. The first flexible coupling and the second flexible coupling are subject to, with respect to the engine central axis, torsional motion and lateral motion. The first and second flexible couplings individually having a torsional stiffness under the torsional motion and a lateral stiffness under the lateral motion, and a ratio, of each of the first and second flexible couplings, of the torsional stiffness to the lateral stiffness is greater than or equal to 2. 1. A gas turbine engine comprising:a fan;a fan shaft coupled with the fan and arranged along an engine central axis;a gear system rotatably coupled to the fan shaft; andfirst and second flexible couplings supporting the gear system, the first flexible coupling and the second flexible coupling being subject to, with respect to the engine central axis, torsional motion and lateral motion,the first and second flexible couplings individually having a torsional stiffness under the torsional motion and a lateral stiffness under the lateral motion, and a ratio, of each of the first and second flexible couplings, of the torsional stiffness to the lateral stiffness is greater than or equal to 2.2. The gas turbine engine as recited in claim 1 , wherein the gear system includes a sun gear in meshed engagement with multiple intermediate gears that are rotatably mounted on bearings in a non-rotatable carrier claim 1 , each intermediate gear is in meshed engagement with a rotatable ring gear claim 1 , the sun gear is rotatably coupled to the fan shaft claim 1 , and the first flexible coupling is coupled with the non-rotatable carrier.3. The gas turbine engine as recited in claim 2 , wherein the gear system is coupled through an input shaft to a low pressure turbine claim 2 , the low pressure turbine ...

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Номер записи: 111
31-01-2019 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20190032572A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a non-rotatable flexible coupling having Stiffnesses A, B, C, D and E, which are defined herein, and at least ratio of a frame lateral stiffness FLS as follows: FLS/Stiffness A in a range of about 6.0 to about 25.0, FLS/Stiffness B in a range of about 10.0 to about 40.0, FLS/Stiffness C in a range of about 1.5 to about 7.0, FLS/Stiffness D in a range of about 0.25 to about 0.50, or FLS/Stiffness E in a range of about 6.0 to about 40.0. 1. A gas turbine engine comprising:a fan;a fan shaft arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a frame lateral stiffness (FLS);an input shaft;a high pressure turbine and a low pressure turbine, the low pressure turbine having a pressure ratio of greater than 5;a gear system rotatably coupled to the fan shaft, the gear system having a gear reduction ratio that is greater than 2.3; anda non-rotatable flexible coupling at least partially supporting the gear system, the non-rotatable flexible coupling being subject to at least one type of motion, with respect to the engine central axis, selected from the group consisting of Motion I, Motion II, Motion III, Motion IV and combinations thereof, wherein, Motion I is parallel offset guided end motion, Motion II is cantilever beam free end motion, Motion III is angular misalignment no offset motion, and Motion IV is axial motion,the non-rotatable flexible coupling having Stiffnesses A, B, C, D and E, wherein Stiffness A is axial stiffness under Motion IV, Stiffness B is radial stiffness under Motion II, Stiffness C is radial stiffness under Motion I, Stiffness D is torsional stiffness under Motion I, and Stiffness E is angular stiffness under Motion III, and at least one ofa ratio of FLS/Stiffness A of the non-rotatable flexible coupling is in a range of about 6.0 to about 25.0,a ratio of FLS/Stiffness B of the non-rotatable flexible coupling is in a range of about 10.0 to about 40.0,a ratio of FLS/Stiffness C of the ...

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Номер записи: 112
30-01-2020 дата публикации

TURBOFAN ENGINE FRONT SECTION

Номер: US20200032715A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between 219 and 328. 1. A turbofan engine comprising:a fan section including a plurality of fan blades;a compressor section including a compressor entrance passage including an inlet disposed at an inlet diameter and an outlet to the compressor disposed at an outlet diameter, wherein a ratio of the inlet diameter to the outlet diameter is between 1.22 and 1.82;a combustor receiving airflow from the compressor section and generating a high-energy flow;a turbine section driven by the high-energy flow;a geared architecture configured to be driven by the turbine section for rotating the fan hub at a speed different than the turbine section, wherein a geared architecture power transfer parameter (PTP) is defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between 219 and 328.2. The turbofan engine as recited in claim 1 , wherein the geared architecture includes an axial length between 3.03 and 4.60 inches.3. The turbofan engine as recited in claim 2 , wherein the turbofan engine includes an overall axial distance from a forward part of the fan hub to a forward bearing assembly and a ratio of the overall axial distance to the axial length of the geared architecture is between 6 and 18.4. The turbofan engine as recited in claim 2 , wherein the geared architecture comprises an epicyclic gear system including a ring gear circumscribing a plurality of planetary gears driven by a sun gear and a carrier ...

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Номер записи: 113
18-02-2016 дата публикации

Turbofan Engine Main Bearing Arrangement

Номер: US20160047273A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine () comprises a fan (), a fan drive gear system (), a fan shaft () coupling the fan drive gear system to the fan, a low spool, an intermediate spool, and a core spool. The low spool engages at least three main bearings of which at least two are non-thrust bearings and at least one is a thrust bearing. The fan shaft engages at least two bearings (). The core spool engages at least two bearings (). The intermediate spool engages at least two of said bearings (--). 2. The engine of claim 1 , wherein:said thrust bearing is a non-duplex ball bearing.3. The engine of claim 1 , wherein:said non-thrust bearings and said thrust bearing are rolling element bearings.4. The engine of claim 3 , wherein:said non-thrust bearings are roller bearings; andsaid thrust bearing is a ball bearing.5. The engine of claim 1 , wherein:there are exactly nine said main bearings.6. The engine of claim 5 , wherein:each of said nine said main bearings is either a single stage rolling element bearing or a multi-stage bearing wherein an interstage gap is no more than 30 mm.7. The engine of claim 5 , wherein:the at least two non-thrust bearings engaging the low spool are exactly two.8. The engine of claim 7 , wherein:said one thrust bearing engaging the low spool also engages the intermediate shaft.9. The engine of claim 7 , wherein:{'b': '22', 'said two non-thrust bearings engaging the low spool also engage a case ().'}10. The engine of claim 5 , wherein:of the at least two bearings that engage the fan shaft, at least one is a non-thrust bearing and at least one is a thrust bearing.11. The engine of claim 5 , wherein:of the at least two bearings that engage the core spool, at least one is a non-thrust bearing and at least one is a thrust bearing.12. The engine of claim 5 , wherein:of the at least two bearings that engage the intermediate spool, at least one is a non-thrust bearing and at least one is a thrust bearing.13. The engine of claim 1 , wherein:of the at least two bearings ...

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Номер записи: 114
15-02-2018 дата публикации

GEARED TURBOFAN WITH LOW SPOOL POWER EXTRACTION

Номер: US20180045119A1
Автор: McCune Michael E., Sheridan William G., Suciu Gabriel L.
Принадлежит:

A geared turbofan engine includes a first spool including a first compressor and a first turbine. The first compressor is immediately before a combustor and the first turbine is immediately after the combustor. A second spool includes at least a second turbine disposed axially aft of the first turbine. A first tower shaft is engaged to drive the high speed spool. A second tower shaft engaged to be driven by the second spool. A starter is engaged to drive the first tower shaft. An accessory gear box is driven by the second tower shaft. A first clutch is disposed between the first tower shaft and the starter. The first clutch is configured to enable the starter to drive the high speed spool. A second clutch is disposed between the second tower shaft and the accessory gear box, the second clutch configured to enable the second spool to drive the accessory gear box. A gas turbine engine and a method of operating a gas turbine engine are also disclosed. 1. A geared turbofan engine comprising:a first spool including a first compressor and a first turbine, wherein the first compressor is immediately before a combustor and the first turbine is immediately after the combustor;a second spool including at least a second turbine disposed axially aft of the first turbine;a first tower shaft engaged to drive the first spool;a second tower shaft engaged to be driven by the second spool;an accessory gear box coupled to both the first tower shaft and the second tower shaft;a starter engaged to drive the accessory gear box;a first clutch disposed between the first tower shaft and the accessory gear box, the first clutch configured to enable the starter to drive the first spool through the accessory gear box, wherein the first clutch is a one-way mechanical clutch capable of transmitting torque only from the accessory gear box to the first tower shaft; anda second clutch disposed between the second tower shaft and the accessory gear box, the second clutch configured to enable the ...

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Номер записи: 115
19-02-2015 дата публикации

OIL BAFFLE FOR GAS TURBINE FAN DRIVE GEAR SYSTEM

Номер: US20150051040A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

An epicyclic gear train component includes spaced apart walls with circumferentially spaced mounts that interconnect the walls. The mounts provide circumferentially spaced apart apertures between the mounts at an outer circumference of the walls. Baffles are arranged between the walls near the mounts. Gear pockets are provided between the baffles and the baffles include a lubrication passage that terminates at least one of the gear pockets. One of the walls includes a hole which has a tube that extends through the hole and is received in an opening in the baffle. The tube is in communication with the lubrication passage. 1. An epicyclic gear train component comprising:spaced apart walls with circumferentially spaced mounts interconnecting the walls, the mounts providing circumferentially spaced apart apertures between the mounts at an outer circumference of the walls;baffles arranged between the walls near the mounts,gear pockets provided between the baffles, and the baffles including a lubrication passage terminating at least one of the gear pockets; andwherein one of the walls includes a hole, and a tube extends through the hole and is received in an opening in the baffle, the tube in communication with the lubrication passage.2. The epicyclic gear train component according to claim 1 , wherein the baffles are secured to at least one of the walls and the mounts by a fastening element.3. An epicyclic gear train component comprising:spaced apart walls with circumferentially spaced mounts interconnecting the walls, the mounts providing circumferentially spaced apart apertures between the mounts at an outer circumference; andbaffles arranged between the walls near the mounts,gear pockets provided between the baffles, and the baffles including a lubrication passage terminating at least one of the gear pockets,wherein the lubrication passage includes a primary passage extending laterally between the walls and first and second passages in communication with the primary ...

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Номер записи: 116
10-03-2022 дата публикации

GAS TURBINE ENGINE FRONT SECTION

Номер: US20220074350A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio. 1. A gas turbine engine comprising:a propulsor section;a compressor section including a compressor entrance passage including an inlet disposed at an inlet diameter and an outlet to the compressor disposed at an outlet diameter,;a combustor receiving airflow from the compressor section and generating a high-energy flow;a turbine section driven by the high-energy flow;a geared architecture configured to be driven by the turbine section for driving the propulsor section at a speed different than the turbine section, wherein a geared architecture power transfer parameter (PTP) is defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between 219 and 328.2. The gas turbine engine as recited in claim 1 , wherein the geared architecture includes an axial length between 3.03 and 4.60 inches.3. The gas turbine engine as recited in claim 2 , wherein the gas turbine engine includes an overall axial distance from a forward part of the propulsor section to a forward bearing assembly and a ratio of the overall axial distance to the axial length of the geared architecture is between 6 and 18.4. The gas turbine engine as recited in claim 2 , wherein the geared architecture comprises an epicyclic gear system including a ring gear circumscribing a plurality of planetary gears driven by a sun gear and a carrier supporting the planetary gears and the gear volume is defined within a space bounded by the ring gear and ...

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Номер записи: 117
10-03-2022 дата публикации

GAS TURBINE ENGINE WITH NON-EPICYCLIC GEAR REDUCTION SYSTEM

Номер: US20220074351A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan drive turbine driving a shaft. The shaft engages a gear reduction. The gear reduction drives a fan rotor at a speed that is less than the speed of the fan drive turbine. The gear reduction is a non-epicyclic gear reduction. 1. A gas turbine engine comprising:a propulsor drive turbine driving a turbine shaft, said turbine shaft engaging a gear reduction, said gear reduction driving a propulsor rotor at a speed that is less than the speed of the propulsor drive turbine;said gear reduction being a non-epicyclic gear reduction;wherein the propulsor rotor has a propulsor shaft, and the propulsor shaft is co-axial with the propulsor drive turbine; anda pressure ratio across said propulsor drive turbine being greater than 5 to 1, with the pressure ratio defined by a pressure measured prior to an inlet to said propulsor drive turbine related to a pressure at an outlet of said propulsor drive turbine prior to any exhaust nozzle.2. The gas turbine engine as set forth in claim 1 , wherein said turbine shaft drives an input gear of said gear reduction through a flexible input shaft.3. The gas turbine engine as set forth in claim 2 , wherein said propulsor shaft is a flexible output shaft.4. The gas turbine engine as set forth in claim 3 , wherein said input gear drives a plurality of circumferentially spaced intermediate gears to claim 3 , in turn claim 3 , drive an output gear.5. The gas turbine engine as set forth in claim 4 , wherein said intermediate gears have a first set of gear teeth engaging gear teeth on said input gear and a second set of gear teeth engaging gear teeth on said output gear.6. The gas turbine engine as set forth in claim 5 , wherein there are a plurality of oil baffles circumferentially spaced between said intermediate gears and for delivering a lubricant to said input gear.7. The gas turbine engine as set forth in claim 2 , wherein said input gear drives a plurality of circumferentially spaced intermediate gears to ...

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Номер записи: 118
10-03-2022 дата публикации

METHOD OF ASSEMBLY FOR GEAR SYSTEM WITH ROTATING CARRIER

Номер: US20220074353A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A method of assembling a gear system for a gas turbine engine according to an example of the present disclosure includes the steps of providing a unitary carrier, positioning baffles between side walls of the carrier such that an end of each of the baffles abuts a respective mount of the carrier, inserting intermediate gears through a central opening of the carrier, and then moving the intermediate gears radially outwardly into respective gear pockets of the carrier, inserting a sun gear through the central opening, moving the intermediate gears radially inwardly to engage the sun gear, coupling a shaft to the carrier, including attaching a torque frame to the carrier such that the torque frame interconnects the carrier and the shaft, placing first and second ring gear halves of a ring gear onto an outer periphery of the intermediate gears, and fixedly attaching the ring gear to an engine static structure. A gear system is also disclosed. 1. A method of assembling a gear system for a gas turbine engine comprising the steps of:providing a unitary carrier defining a central axis, wherein the unitary carrier includes a pair of axially spaced apart side walls and circumferentially spaced mounts, at least one of the walls defines a central opening, the mounts are circumferentially spaced about the carrier at a distance that is less than a width of a sun gear and a plurality of intermediate gears to establish circumferentially spaced apart peripheral openings at an outer circumference of the carrier, and curved walls of adjacent pairs of the mounts establish gear pockets extending radially from the central opening to the respective peripheral openings at the outer circumference of the carrier relative to the central axis, and;positioning baffles between the side walls of the unitary carrier such that an end of each of the baffles abuts a respective one of mounts, wherein each of the baffles includes a curved surface that defines a respective one of the gear pockets; ...

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Номер записи: 119
20-02-2020 дата публикации

ROLLER BEARINGS FOR HIGH RATIO GEARED TURBOFAN ENGINE

Номер: US20200056498A1
Автор: Sheridan William G.
Принадлежит:

A gear system for a geared turbofan engine is disclosed. The gear system includes a sun gear driven by a low spool shaft. The sun gear defines a sun gear diameter. A rotating carrier drives a fan. The carrier defines an outer carrier diameter and an inner carrier diameter. A non-rotating ring gear is also included. The ring gear defines a ring gear diameter and the ring gear diameter is smaller than the outer carrier diameter. A set of planet gears are mounted on corresponding rolling element bearing assemblies. Each roller element bearing assembly is supported within the carrier within a space defined between the carrier outer diameter and the carrier inner diameter. Each of the sun gear, ring gear and planet gears are substantially centered along a gearbox centerline transverse to an engine longitudinal axis and the gear system provides a speed reduction ratio between an input to the sun gear and an output from the carrier between 3:1 and 5:1. A method of creating a gear system for a geared turbofan engine and a geared turbofan system are also disclosed. 1. A gear system for a geared turbofan engine , the gear system comprising:a sun gear driven by a low spool shaft, the sun gear defining a sun gear diameter;a rotating carrier that drives a fan, the carrier defining an outer carrier diameter and an inner carrier diameter;a non-rotating ring gear, the ring gear defining a ring gear diameter, the ring gear diameter is smaller than the outer carrier diameter; anda set of planet gears including 3, 4 or 5 planet gears; androlling element bearing assemblies supporting each of the set of planet gears, each roller element bearing assembly is supported within the carrier within a space defined between the carrier outer diameter and the carrier inner diameter, wherein a diameter of each of the bearing elements of the roller element bearing assembly is less than a sum of the carrier outer diameter and the carrier inner diameter multiplied by √ and divided by twice the number ...

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Номер записи: 120
20-02-2020 дата публикации

AUXILIARY OIL SYSTEM FOR GEARED GAS TURBINE ENGINE

Номер: US20200056544A1
Автор: Sheridan William G.
Принадлежит:

A gas turbine engine comprises a fan drive turbine, a fan rotor, and a gear reduction driven by the fan drive turbine to, in turn, drive the gear architecture. A main oil supply system supplies oil to components within the gear reduction, and an auxiliary oil supply system. The auxiliary oil system operates to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail. 1. A gas turbine engine comprising:a fan drive turbine, a fan rotor, and a gear reduction driven by said fan drive turbine to, in turn, drive said gear architecture, a main oil supply system for supplying oil to components within said gear reduction, and an auxiliary oil supply system;said gear reduction includes a sun gear being driven by said fan drive turbine to drive intermediate gears that engage a ring gear;said auxiliary oil system being operable to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail;said auxiliary oil system being operable to allow the engine to operate under windmill conditions in the air for up to 90 minutessaid auxiliary oil system being operable to operate indefinitely on the ground when windmilling with wind speeds below 85 mph or less; andsaid auxiliary oil system being operable to fly with the engine in an aircraft under negative gravity conditions for at least 20 seconds.2. The gas turbine engine as set forth in claim 1 , wherein said sun gear claim 1 , said intermediate gears and said ring gear are enclosed in a bearing compartment claim 1 , which captures oil removed via a scavenge line connected to a main oil pump.3. The gas turbine engine as set forth in claim 2 , wherein said main oil pump has a scavenge stage that directs scavenged oil to a main oil tank.4. The gas turbine engine as set forth in claim 3 , wherein oil in said main oil tank feeds a main pump pressure stage which then ...

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Номер записи: 121
22-05-2014 дата публикации

OIL SYSTEM BEARING COMPARTMENT ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20140140824A1
Автор: Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine with a geared architecture includes a multiple of bearing compartments and at least one carbon seal that seals at least one side of each of the multiple of bearing compartments. 1. A gas turbine engine with a geared architecture , comprising:a multiple of bearing compartments; andat least one carbon seal which seals at least one side of each of said multiple of bearing compartments.2. The gas turbine engine as recited in claim 1 , wherein said multiple of bearing compartments include a front bearing compartment.3. The gas turbine engine as recited in claim 1 , wherein said multiple of bearing compartments include a mid bearing compartment.4. The gas turbine engine as recited in claim 1 , wherein said multiple of bearing compartments include a mid-turbine bearing compartment.5. The gas turbine engine as recited in claim 1 , wherein said multiple of bearing compartments include a rear bearing compartment.6. The gas turbine engine as recited in claim 1 , wherein each of said multiple of bearing compartments interface with an engine shaft.7. The gas turbine engine as recited in claim 1 , wherein said geared architecture is 98% efficient.8. The gas turbine engine as recited in claim 1 , wherein said multiple of bearing compartments include:a front bearing compartment;a mid bearing compartment axially aft of said front bearing compartment;a mid-turbine bearing compartment axially aft of said mid bearing compartment; anda rear bearing compartment axially aft of said mid-turbine bearing compartment.9. A gas turbine engine claim 1 , comprising:a front bearing compartment bounded by a first and second carbon seal;a mid bearing compartment bounded by a first and second carbon seal, said mid bearing compartment axially aft of said front bearing compartment;a mid-turbine bearing compartment bounded by a first and second carbon seal, said mid-turbine bearing compartment axially aft of said mid bearing compartment; anda rear bearing compartment bounded by a ...

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Номер записи: 122
28-02-2019 дата публикации

Geared gas turbine engine with oil deaerator

Номер: US20190063330A1
Автор: William G. Sheridan
Принадлежит: United Technologies Corp

A gas turbine engine includes a fan section including a fan rotor, a compressor section, a turbine section including a fan drive turbine that drives the fan rotor through a gear reduction, and a lubrication system that supplies oil to the gear reduction and includes a lubricant pump that supplies a mixed air and oil to an inlet of a deaerator during operation. The deaerator includes a shell defining a cavity and a separator that separates the mixed air and oil in the cavity, delivers separated air to an air outlet of the deaerator and delivers separated oil back into an oil tank during operation. A method of designing a gas turbine engine is also disclosed.

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Номер записи: 123
10-03-2016 дата публикации

TURBOFAN ENGINE FRONT SECTION

Номер: US20160069270A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between about 219 and 328. 1. A turbofan engine comprising:a fan section including a fan hub including a hub diameter supporting a plurality of fan blades including a tip diameter with a ratio of the hub diameter to the tip diameter is between about 0.24 and about 0.36;a compressor section;a combustor receiving airflow from the compressor section and generating a high-energy flow;a turbine section driven by the high-energy flow; and{'sup': 3', '3, 'a geared architecture driven by the turbine section for rotating the fan hub at a speed different than the turbine section, the geared architecture including a gear volume between about 526 inand about 790 in.'}2. The turbofan engine as recited in claim 1 , wherein the geared architecture includes an axial length between about 3.03 and about 4.60 inches.3. The turbofan engine as recited in claim 2 , wherein the turbofan engine includes an overall axial distance from a forward part of the fan hub to a forward bearing assembly and a ratio of the overall axial distance to the axial length of the geared architecture is between about 6 and about 18.4. The turbofan engine as recited in claim 2 , wherein the geared architecture comprises an epicyclic gear system including a ring gear circumscribing a plurality of planetary gears driven by a sun gear and a carrier supporting the planetary gears and the gear volume is defined within a space bounded by the ring gear and outer periphery of the carrier.5. The turbofan engine as recited ...

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Номер записи: 124
08-03-2018 дата публикации

METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE

Номер: US20180066590A1
Автор: Hasel Karl L., Sheridan William G.
Принадлежит:

In one exemplary embodiment, a gas turbine engine includes a fan section having a fan with a low corrected fan tip speed less than 1400 ft/sec. A bypass ratio is greater than 11.0 and less than 22.0. A fan pressure ratio is less than 1.48. A speed reduction device includes a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1. A low pressure turbine including three stages and a high pressure turbine including two stages. The low pressure turbine includes at least one rotor constrained by a first stress level. At least one of a plurality of fan blades is constrained by a second stress level and has a fan tip speed boundary condition. The gear ratio is configured such that the at least one fan blade does not exceed the fan tip speed boundary condition or the second stress level. 1. A gas turbine engine comprising:an engine centerline longitudinal axis;{'b': '1400', 'sup': '0.5', 'a fan section including a fan with a plurality of fan blades and rotatable about the engine centerline longitudinal axis, wherein the fan has a low corrected fan tip speed less than ft/sec, wherein the low corrected fan tip speed is an actual fan tip speed at an ambient temperature divided by [(Tram° R)/(518.7° R)], where Tram represents the ambient temperature in degrees Rankine;'}a bypass ratio greater than 11.0 and less than 22.0;a fan pressure ratio less than 1.48, wherein the fan pressure ratio is measured across a fan blade alone;a speed reduction device comprising a gear system with a gear ratio of at least 2.6 and less than or equal to 4.1;a low pressure turbine in communication with a first shaft;a high pressure turbine in communication with a second shaft;wherein the first shaft and second shaft are concentric and mounted via at least one of a plurality of bearing systems for rotation about the engine centerline longitudinal axis, and the first shaft is in communication with the fan through the speed reduction device;wherein the low pressure turbine includes ...

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Номер записи: 125
27-02-2020 дата публикации

TURBOFAN WITH MOTORIZED ROTATING INLET GUIDE VANE

Номер: US20200063657A1
Автор: Kupratis Daniel Bernard, Sheridan William G., Sidelkovskiy Dmitriy B., Terwilliger Neil
Принадлежит:

A gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible things, a fan including a plurality of fan blades rotatable about an axis, a plurality of inlet guide vanes mounted forward of the plurality of fan blades, the plurality of inlet guide vanes selectively rotatable about the axis independent of the plurality of fan blades; and a motor for controlling rotation of the plurality of inlet guide vanes about the axis. 1. A gas turbine engine comprising:a fan including a plurality of fan blades rotatable about an engine central longitudinal axis;a plurality of inlet guide vanes mounted forward of the plurality of fan blades, the plurality of inlet guide vanes selectively rotatable about the engine central longitudinal axis independent of the plurality of fan blades;a geared architecture coupled to the fan;at least one fan exit guide vane mounted aft of the plurality of fan blades;a motor for controlling rotation of the plurality of inlet guide vanes about the engine central longitudinal axis; andan electric conduit in electrical communication with the motor, the electrical conduit extending through the at least one fan exit guide vane and the geared architecture .2. The gas turbine engine as recited in claim 1 , wherein the plurality of inlet guide vanes are movable from a non-rotating condition to a rotating condition independent of rotation of the plurality of fan blades.3. The gas turbine engine as recited in claim 1 , wherein the plurality of inlet guide vanes are rotatable at a speed different than the plurality of fan blades.4. The gas turbine engine as recited in claim 1 , including a fan hub supporting rotation of the plurality of fan blades and the plurality of inlet guide vanes are rotatable supported by the fan hub for rotation separate from the fan hub.5. The gas turbine engine as recited in claim 4 , wherein the motor is supported within the fan hub.6. The gas turbine engine as recited in claim 1 , wherein ...

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Номер записи: 126
17-03-2016 дата публикации

FAN DRIVE GEAR SYSTEM WITH IMPROVED MISALIGNMENT CAPABILITY

Номер: US20160076393A1
Автор: Sheridan William G.
Принадлежит:

An epicyclic gear assembly according to an exemplary aspect of the present disclosure includes, 60 among other things, a carrier including a first plate axially spaced from a second plate by a radially outer connector. A first set of epicyclic gears supported adjacent the first plate include a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear. A second set of epicyclic gears are axially spaced from the first set of epicyclic gears and supported adjacent the second plate, and include a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear. The first epicyclic gear set and the second epicyclic gear set maintain relative intermeshing alignment during flexure induced deformation of the carrier. 1. An epicyclic gear assembly comprising:a carrier including a first plate axially spaced from a second plate by a radially outer connector;a first set of epicyclic gears supported adjacent the first plate, including a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear; anda second set of epicyclic gears axially spaced from the first set of epicyclic gears and supported adjacent the second plate, and including a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear, whereby the first epicyclic gear set and the second epicyclic gear set maintain relative intermeshing alignment during flexure induced deformation of the carrier.2. The assembly of claim 1 , including at least one intermediate gear opening in the carrier claim 1 , wherein one of the first set of intermediate gears and one of the second set of intermediate gears are located within the at least one intermediate gear opening.3. The assembly of claim 1 , wherein the first set of intermediate gears move independently of the second set of intermediate gears.4. The assembly of claim 1 , wherein ...

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Номер записи: 127
15-03-2018 дата публикации

A FAN DRIVE GEAR SYSTEM FOR DRIVING A FAN IN A GAS TURBINE ENGINE HAVING A HIGH BYPASS RATIO

Номер: US20180073439A9
Автор: Hasel Karl L., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section that includes a fan rotatable about an axis of rotation of the gas turbine engine. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an axis of rotation of the gas turbine engine; anda speed reduction device connected to the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5, wherein a bypass ratio is greater than about 11.0.2. The gas turbine engine of claim 1 , wherein the speed reduction device includes a star gear system gear ratio of at least 2.6.3. The gas turbine engine of claim 2 , wherein the speed reduction device includes a system gear ratio less than or equal to 4.1.4. The gas turbine engine of claim 3 , including the bypass ratio is less than about 22.0.5. The gas turbine engine of claim 4 , wherein the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.6. The gas turbine engine of claim 1 , wherein the star system includes a sun gear claim 1 , a plurality of star gears claim 1 , a ring gear claim 1 , and a carrier.7. The gas turbine engine of claim 6 , wherein each of the plurality of star gears includes at least one bearing.8. The gas turbine engine of claim 7 , wherein the carrier is fixed from rotation.9. The gas turbine engine of claim 8 , wherein a low pressure turbine is mechanically attached to the sun gear.10. The gas turbine engine of claim 9 , wherein a fan section is mechanically attached to the ring gear.11. The gas turbine engine of claim 1 , wherein an input of the speed reduction device is rotatable in a first direction and an output of the speed reduction device is rotatable in a second direction opposite to the first direction.12. The gas turbine engine of ...

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Номер записи: 128
07-03-2019 дата публикации

FAN DRIVE GEAR SYSTEM

Номер: US20190072040A1
Автор: McCune Michael E., Roberge Gary D., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. The speed change mechanism is an epicyclic gear train. A torque frame surrounds the speed change mechanism and includes a plurality of fingers. A bearing support is attached to the plurality of fingers. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted on the bearing support aft of the speed change mechanism. The second fan section bearing is a fan thrust bearing. 1. A gas turbine engine comprising:a fan section;a speed change mechanism for driving the fan section, wherein the speed change mechanism is an epicyclic gear train;a torque frame surrounds the speed change mechanism and includes a plurality of fingers;a bearing support attached to the plurality of fingers; anda first fan section support bearing mounted forward of the speed change mechanism and a second fan section bearing mounted on the bearing support aft of the speed change mechanism, wherein the second fan section bearing is a fan thrust bearing.2. The gas turbine engine of claim 1 , wherein the fan thrust bearing engages a gas turbine static structure and the bearing support.3. The gas turbine engine of claim 2 , wherein the bearing support is attached to distal ends of the plurality of fingers.4. The gas turbine engine of claim 3 , wherein the epicyclic gear train is a planetary gear system including a sun gear in communication with a fan drive turbine and a planet carrier in communication with the fan section.5. The gas turbine engine of claim 4 , wherein the torque frame includes a first end for engaging the fan section and second end supporting the second fan section bearing.6. The gas turbine engine of claim 5 , wherein each of the plurality of fingers include at least one groove.7. The gas turbine engine of claim 6 , wherein the bearing support includes a plurality of tangs that engage a corresponding one of the at ...

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Номер записи: 129
05-06-2014 дата публикации

OIL BAFFLE FOR GAS TURBINE FAN DRIVE GEAR SYSTEM

Номер: US20140154054A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbine engine comprises compressor and turbine sections. An epicyclic gear train includes a carrier, a sun gear and intermediate gears arranged about and intermeshing with the sun gear. The intermediate gears are supported by the carrier. A baffle includes a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on at least one of the sun gear and intermediate gears. A method of designing a turbine engine is also disclosed. 1. A turbine engine comprising:compressor and turbine sections; and a carrier;', 'a sun gear and intermediate gears arranged about and intermeshing with the sun gear, the intermediate gears supported by the carrier; and', 'a baffle including a lubrication passage near at least one of the sun gear and intermediate gears for directing a lubricant on the at least one of the sun gear and intermediate gears., 'an epicyclic gear train including2. The turbine engine according to claim 1 , comprising a ring gear intermeshing with the intermediate gears and an output shaft interconnected to the ring gear claim 1 , and an input shaft interconnected to the sun gear.3. The turbine engine according to claim 2 , wherein the carrier is fixed relative to a housing claim 2 , the output shaft drives a fan claim 2 , and the input shaft supports a compressor hub having compressor blades.4. The turbine engine according to claim 2 , wherein there are three turbine rotors in the turbine section claim 2 , with a fan drive turbine driving the fan through the input shaft claim 2 , and two other turbine rotors driving compressor rotors.5. The turbine engine according to claim 2 , wherein the carrier is fixed relative to a housing claim 2 , the output shaft drives a compressor hub claim 2 , and a fan at a common speed claim 2 , and the input shaft is driven by a turbine rotor.6. The turbine engine according to claim 5 , wherein there are two turbine rotors claim 5 , with the first turbine rotor driving a second compressor ...

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Номер записи: 130
16-03-2017 дата публикации

GEARED TURBOFAN ENGINE GEARBOX ARRANGEMENT

Номер: US20170074165A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbofan engine according to an exemplary aspect of the present disclosure includes, among other things, a fan having a plurality of blades, and a transmission is configured to drive the fan. The fan blades have a peak tip radius R. The fan blades have an inboard leading edge radius Rat an inboard boundary of the flowpath. A ratio of Rto Ris less than about 0.40. 2. The engine of claim 1 , wherein:each fan blade has a leading edge and a trailing edge;a splitter is positioned along the flowpath through the engine and having a leading rim separates a core branch of the flowpath from a bypass branch of the flowpath;{'sub': 'II', 'an inboard boundary of the core branch of the flowpath has a radius Rat an axial position of the splitter rim;'}{'sub': 'I', 'the inboard boundary of the core branch of the flowpath has a radius Rat a leading stage of blades of the first compressor; and'}{'sub': 10', 'II, 'a ratio of an axial length Lbetween the splitter rim and the leading stage of blades of the first compressor at the inboard boundary of the core branch of the flowpath to the radius Ris less than 1.2.'}3. The engine of claim 1 , wherein:each fan blade has a leading edge and a trailing edge;a splitter is positioned along the flowpath through the engine and having a leading rim separates a core branch of the flowpath from a bypass branch of the flowpath;{'sub': 'II', 'an inboard boundary of the core branch of the flowpath has a radius Rat an axial position of the splitter rim;'}{'sub': 'I', 'the inboard boundary of the core branch of the flowpath has a radius Rat a leading stage of blades of the first compressor; and'}{'sub': I', 'II, 'a ratio of the radius Rto the radius Ris greater than 0.50 and is less than 1.0.'}4. The engine of claim 3 , wherein:{'sub': I', 'II, 'said ratio of the radius Rto the radius Ris 0.55-1.0.'}5. The engine of claim 4 , wherein:the fan blades are non-variable.6. The engine of claim 5 , further comprising:a variable fan nozzle.7. The engine of ...

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Номер записи: 131
05-06-2014 дата публикации

GEARED ARCHITECTURE GAS TURBINE ENGINE WITH IMPROVED LUBRICATION AND MISALIGNMENT TOLERANT ROLLER BEARING SYSTEM

Номер: US20140155213A1
Автор: Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine includes a structural oil baffle housing which at least partially supports a set of intermediate gears. 1. A gas turbine engine , comprising:a geared architecture with a multiple of intermediate gears; anda structural oil baffle housing that at least partially supports said set of intermediate gears.2. The gas turbine engine as recited in claim 1 , wherein each of said multiple of intermediate gears is mounted to a respective flexible carrier post.3. The gas turbine engine as recited in claim 2 , further comprising a spherical joint mounted to each flexible carrier post.4. The gas turbine engine as recited in claim 2 , wherein said structural oil baffle housing is mounted to a spherical joint mounted to each flexible carrier post.5. The gas turbine engine as recited in claim 1 , further comprising a rotationally fixed carrier claim 1 , each of said multiple of intermediate gears mounted to a respective flexible carrier post which extends from said carrier.6. The gas turbine engine as recited in claim 5 , further comprising an oil manifold defined by said carrier.7. The gas turbine engine as recited in claim 6 , wherein said oil manifold includes a first oil circuit and a second oil circuit.8. The gas turbine engine as recited in claim 7 , wherein said first oil circuit communicates with each respective flexible carrier post.9. The gas turbine engine as recited in claim 7 , wherein said second oil circuit communicates with a multiple of oil nozzles.10. The gas turbine engine as recited in claim 1 , further comprising a rolling element bearing mounted between said structural oil baffle housing and each of said multiple of intermediate gears.11. The gas turbine engine as recited in claim 10 , wherein said structural oil baffle housing directs oil to said multiple of intermediate gears.12. The gas turbine engine as recited in claim 1 , wherein said geared architecture includes a planetary gear system.13. The gas turbine engine as recited in claim 1 ...

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Номер записи: 132
24-03-2016 дата публикации

FAN DRIVE GEAR SYSTEM

Номер: US20160084104A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted aft of the speed change mechanism. 1. A gas turbine engine comprising:a fan section;a speed change mechanism for driving the fan section; anda first fan section support bearing mounted forward of the speed change mechanism and a second fan section bearing mounted aft of the speed change mechanism.2. The gas turbine engine of claim 1 , wherein the speed change mechanism is a planetary gear system including a sun gear in communication with a fan drive turbine and a planet carrier in communication with the fan section.3. The gas turbine engine of claim 1 , wherein a torque frame surrounds the speed change mechanism.4. The gas turbine engine of claim 3 , wherein the torque frame includes a first end for engaging the fan section and second end supporting the second fan section bearing.5. The gas turbine engine of claim 4 , wherein the torque frame includes a plurality of fingers that surround a planet carrier of the speed change mechanism.6. The gas turbine engine of claim 5 , wherein the second fan section bearing is a fan thrust bearing supported on a bearing support attached to distal ends of the plurality of fingers on the torque frame.7. The gas turbine engine of claim 6 , wherein the bearing support includes at least one tang that engages a groove in at least one of the plurality of fingers.8. The gas turbine engine of claim 7 , wherein the groove is located on a radially inner side of at least one of the plurality of fingers.9. The gas turbine engine of claim 1 , wherein the speed change mechanism is at least partially axially aligned with a compressor section.10. The gas turbine engine of claim 1 , further comprising a high pressure compressor with a compression ratio of approximately 20:1 or greater.11. The gas turbine engine of claim ...

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Номер записи: 133
24-03-2016 дата публикации

FAN DRIVE GEAR SYSTEM

Номер: US20160084105A1
Автор: McCune Michael E., Roberge Gary D., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted aft of the speed change mechanism. 1. A gas turbine engine comprising:a fan section;a speed change mechanism for driving the fan section; anda first fan section support bearing mounted forward of the speed change mechanism and a second fan section bearing mounted aft of the speed change mechanism.2. The gas turbine engine of claim 1 , wherein the speed change mechanism is a planetary gear system including a sun gear in communication with a fan drive turbine and a planet carrier in communication with the fan section.3. The gas turbine engine of claim 1 , wherein a torque frame surrounds the speed change mechanism.4. The gas turbine engine of claim 3 , wherein the torque frame includes a first end for engaging the fan section and second end supporting the second fan section bearing.5. The gas turbine engine of claim 4 , wherein the torque frame includes a plurality of fingers that surround a planet carrier of the speed change mechanism.6. The gas turbine engine of claim 5 , wherein the second fan section bearing is a fan thrust bearing supported on a bearing support attached to distal ends of the plurality of fingers on the torque frame.7. The gas turbine engine of claim 6 , wherein the bearing support includes at least one tang that engages a groove in at least one of the plurality of fingers.8. The gas turbine engine of claim 7 , wherein the groove is located on a radially inner side of at least one of the plurality of fingers.9. The gas turbine engine of claim 1 , wherein the speed change mechanism is at least partially axially aligned with a compressor section.10. The gas turbine engine of claim 1 , further comprising a high pressure compressor with a compression ratio of approximately 20:1 or greater and a fan bypass ratio of ...

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Номер записи: 134
25-03-2021 дата публикации

GEARED TURBOFAN WITH LOW PRESSURE TURBINE FEATURES

Номер: US20210087944A1
Автор: Sheridan William G.
Принадлежит:

A turbine engine has a fan and a low pressure compressor that rotate at a common speed and in a common direction. A fan drive turbine drives a gear reduction to, in turn, drive the low pressure compressor and the fan at a speed which is slower than a speed of the fan drive turbine. A combustor intermediate the low pressure compressor and the fan drive turbine and a thrust bearing mount the fan drive turbine, the thrust bearing being aft of a location of the combustor. A shear section in a drive connection connecting the fan drive turbine to the gear reduction is weaker than other portions of the drive connection. The shear section is aft of the thrust bearing. 1. A gas turbine engine comprising:a fan and a low pressure compressor, said fan and said low pressure compressor configured for rotating at a common speed and in a common direction;a fan drive turbine configured for driving a gear reduction, in turn configured for driving said low pressure compressor and said fan at a speed which is slower than a speed of said fan drive turbine;a combustor intermediate said low pressure compressor and said fan drive turbine and a thrust bearing configured for mounting said fan drive turbine, said thrust bearing being aft of a location of said combustor;said fan drive turbine is configured to be supported on said thrust bearing and on a second bearing which is aft of said thrust bearing;a shear section in a drive connection configured for connecting said fan drive turbine to said gear reduction that is weaker than other portions of said drive connection, and said shear section being aft of said thrust bearing;said gear reduction is configured to be supported on a bearing forward of said gear reduction and on a second bearing which is aft of said gear reduction; andsaid fan delivering air into a bypass duct as propulsion air, and delivering air into said low pressure compressor, and a bypass ratio of the air delivered into the bypass duct divided by the volume of air delivered ...

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Номер записи: 135
21-03-2019 дата публикации

PROCESSES FOR TREATMENT OF SPENT ALKALINE WASTE STREAMS

Номер: US20190084854A1
Автор: Lupton F. Stephen, Ranade Bipin R., Sheridan William G.
Принадлежит:

Processes for treatment of spent alkaline stream is disclosed. The process includes passing a spent alkaline stream comprising one or more sulfide compounds and one or more organic compounds to a sulfide oxidation reactor for partial oxidation of the one or more sulfide compounds to provide an effluent stream comprising one or more thiosulfate compounds. The effluent stream is passed to a biological treatment unit to oxidize the one or more thiosulfate compounds to one or more sulfate compounds and biodegrade the one or more organic compounds to carbon dioxide and water to provide a treated alkaline stream. 1. A process for treatment of a spent alkaline stream , comprising:passing the spent alkaline stream comprising one or more sulfide compounds and one or more organic compounds to a sulfide oxidation reactor for partial oxidation of the one or more sulfide compounds to provide an effluent stream comprising one or more thiosulfate compounds; andpassing the effluent stream to a biological treatment unit to oxidize the one or more thiosulfate compounds to one or more sulfate compounds and biodegrade the one or more organic compounds to carbon dioxide and water to provide a treated alkaline stream.2. The process of claim 1 , wherein the spent alkaline stream comprises about 10 to about 170 g/L of sodium sulfide and about 100-10 claim 1 ,000 wppm of mercaptans.3. The process of claim 1 , wherein the sulfide oxidation reactor operates at a temperature of about 75° C. to about 120° C.4. The process of claim 1 , wherein the sulfide oxidation reactor operates at a pressure of about 400 kPa(g) to about 1000 kPa(g).5. The process of further comprising adding an oxygen-containing gas claim 1 , a carbon dioxide stream and a steam stream to the sulfide oxidation reactor for partial oxidation of the one or more sulfide compounds.6. The process of claim 1 , wherein the effluent being passed to biological treatment unit comprises no more than about 100 wppm of sulfides.7. The ...

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Номер записи: 136
21-03-2019 дата публикации

PROVIDING TURBINE ENGINES WITH DIFFERENT THRUST RATINGS

Номер: US20190085771A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A method is provided that includes providing a first turbine engine and providing a second turbine engine. The first turbine engine is configured with a first thrust rating. The first turbine engine includes a first engine rotating assembly and a first engine case structure housing at least the first engine rotating assembly. The second turbine engine is configured with a second thrust rating that is different than the first thrust rating. The second turbine engine includes a second engine rotating assembly and a second engine case structure housing at least the second engine rotating assembly. The first engine case structure and the second engine case structure have at least substantially common configurations. The first turbine engine and the second turbine engine are provided by a common entity. 1. A method , comprising:providing a first turbine engine configured with a first thrust rating, the first turbine engine comprising a first engine rotating assembly and a first engine case structure housing at least the first engine rotating assembly; andproviding a second turbine engine configured with a second thrust rating that is different than the first thrust rating, the second turbine engine comprising a second engine rotating assembly and a second engine case structure housing at least the second engine rotating assembly, wherein the first engine case structure and the second engine case structure have at least substantially common configurations;wherein the first turbine engine and the second turbine engine are provided by a common entity.2. The method of claim 1 , wherein the providing of the first turbine engine comprises assembling the first turbine engine claim 1 , and the providing of the second turbine engine comprises assembling the second turbine engine.3. The method of claim 1 , wherein the first engine case structure and the second engine case structure have identical geometries.4. The method of claim 1 , wherein the first engine case structure and the ...

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Номер записи: 137
05-04-2018 дата публикации

TURBOFAN ENGINE BEARING AND GEARBOX ARRANGEMENT

Номер: US20180094579A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A turbofan engine according to an example of the present disclosure includes, among other things, a fan delivering air into a bypass duct and into a core engine, a fan drive gear system having a gear carrier, and a fan shaft coupling the fan drive gear system to the fan. The core engine includes a low spool including a low pressure turbine and a low pressure shaft, and a high spool including a high pressure turbine, a compressor, and a high pressure shaft coupling the high pressure turbine to the compressor. A first bearing is forward of the fan drive gear system, and supporting said fan shaft and said fan drive gear system, and a second bearing is aft of the fan drive gear system, and supporting the fan drive gear system. 1. A turbofan engine comprising:a fan delivering air into a bypass duct and into a core engine;a fan drive gear system having a gear carrier; anda fan shaft coupling the fan drive gear system to the fan; a low pressure turbine;', 'a low pressure shaft; and, 'the core engine including a low spool comprising a high pressure turbine;', 'a compressor; and', 'a high pressure shaft coupling the high pressure turbine to the compressor;, 'a high spool comprisingwherein a first bearing is forward of the fan drive gear system, and supporting said fan shaft and said fan drive gear system, and a second bearing is aft of the fan drive gear system, and supporting the fan drive gear system.2. The engine as recited in claim 1 , further comprising a case claim 1 , and wherein the fan drive gear system is coupled to the case via a compliant flexure.3. The engine as recited in claim 2 , wherein the second bearing supports the fan shaft relative to a rear support of a static structure claim 2 , the rear support extending inward from a forward frame of the static structure behind the fan drive gear system relative to a longitudinal axis of the core engine.4. The engine as recited in claim 3 , wherein one of the first and second bearings is a thrust bearing claim 3 , ...

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Номер записи: 138
28-03-2019 дата публикации

Deoiler for a gas turbine engine

Номер: US20190093527A1
Автор: William G. Sheridan
Принадлежит: United Technologies Corp

Aspects of the disclosure are directed to a deoiler. The deoiler includes an impeller, a housing arranged as a volute scroll, a splitter wall located at an exit of the deoiler that separates a lubricant and air discharged by the deoiler, and a port located at the exit that discharges the lubricant to at least one of a pump, a tank, or a gearbox.

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Номер записи: 139
26-06-2014 дата публикации

FUNDAMENTAL GEAR SYSTEM ARCHITECTURE

Номер: US20140178180A1
Автор: Sheridan William G.
Принадлежит:

A fan drive gear system for a gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan and a mount flexibly supporting portions of the gear system. A lubrication system supporting the fan drive gear system provides lubricant to the gear system and removes thermal energy produced by the gear system. The lubrication system includes a capacity for removing thermal energy equal to less than about 2% of power input into the gear system. 1. A fan drive gear system for a gas turbine engine comprising:a gear system that provides a speed reduction between a fan drive turbine and a fan; anda lubrication system configured to provide lubricant to the gear system and to remove thermal energy produced by the gear system,wherein the lubrication system includes a capacity for removing thermal energy greater than zero and less than about 2% of power input into the gear system during operation of the engine.2. The fan drive gear system as recited in claim 1 , wherein the gear system transfers power input from the fan drive turbine to the fan at an efficiency greater than about 98% and less than 100%.3. The fan drive gear system as recited in claim 1 , wherein the lubrication system includes a capacity for removing thermal energy equal to less than about 1% of power input into the gear system.4. The fan drive gear system as recited in claim 1 , wherein the lubrication system comprises a main lubrication system configured to provide lubricant to the gear system and an auxiliary lubrication system configured to provide lubricant to the gear system responsive to an interruption of lubricant flow from the main lubrication system.5. The fan drive gear system as recited in claim 1 , wherein the gear system is flexibly supported for movement relative to a static structure of the engine.6. The fan drive gear system as recited in claim 5 , wherein a load limiter is configured to limit movement of the gear system relative to the static ...

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Номер записи: 140
21-04-2016 дата публикации

TURBOFAN ENGINE FRONT SECTION

Номер: US20160108807A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A turbofan engine includes a geared architecture for driving a fan about an axis. The geared architecture includes a sun gear rotatable about an axis, a plurality of planet gears driven by the sun gear and a ring gear circumscribing the plurality of planet gears. A carrier supports the plurality of planet gears. The geared architecture includes a power transfer parameter (PTP) defined as power transferred through the geared architecture divided by gear volume multiplied by a gear reduction ratio and is between about 430 and 645. 1. A turbofan engine comprising:a fan section including a fan hub including a hub diameter supporting a plurality of fan blades including a tip diameter with a ratio of the hub diameter to the tip diameter is between about 0.24 and about 0.36;a compressor section;a combustor receiving airflow from the compressor section and generating a high-energy flow;a turbine section driven by the high-energy flow; and{'sup': 3', '3, 'a geared architecture driven by the turbine section for rotating the fan hub at a speed different than the turbine section, the geared architecture including a gear volume between about 1318 inand about 1977 in.'}2. The turbofan engine as recited in claim 1 , wherein the geared architecture includes an axial length between about 4.16 and about 6.90 inches.3. The turbofan engine as recited in claim 2 , wherein the turbofan engine includes an overall axial distance from a forward part of the fan hub to a forward bearing assembly and a ratio of the overall axial distance to the axial length of the geared architecture is between about 4.8 and about 16.2.4. The turbofan engine as recited in claim 2 , wherein the geared architecture comprises an epicyclic gear system including a ring gear circumscribing a plurality of planetary gears driven by a sun gear and a carrier supporting the planetary gears and the gear volume is defined within a space bounded by the ring gear and outer periphery of the carrier.5. The turbofan engine as ...

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Номер записи: 141
21-04-2016 дата публикации

Turbofan Engine Bearing and Gearbox Arrangement

Номер: US20160108808A1
Автор: Schwarz Frederick M., Sheridan William G., Suciu Gabriel L.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine () comprises a fan (). A fan drive gear system () is configured to drive the fan. A low spool comprises a low pressure turbine () and a low shaft () coupling the low pressure turbine to the fan drive gear system. An intermediate spool comprises an intermediate pressure turbine (), a compressor (), and an intermediate spool shaft () coupling the intermediate pressure turbine to the intermediate spool compressor. A combustor () is between a core spool compressor () and a high pressure turbine (). A first () main bearing engages a static support (′) and a forward hub () of the intermediate spool. A second () main bearing engages the low shaft and the forward hub. 120. A turbofan engine () comprising:{'b': '28', 'a fan ();'}{'b': '60', 'a fan drive gear system () configured to drive the fan;'} [{'b': '50', 'a low pressure turbine (); and'}, {'b': '56', 'a low shaft () coupling the low pressure turbine to the fan drive gear system;'}], 'a low spool comprising [{'b': '48', 'an intermediate pressure turbine ();'}, {'b': '42', 'a compressor (); and'}, {'b': '54', 'an intermediate spool shaft () coupling the intermediate pressure turbine to the intermediate spool compressor;'}], 'an intermediate spool comprising;'} [{'b': '46', 'a high pressure turbine ();'}, {'b': '44', 'a compressor (); and'}, {'b': '52', 'a core shaft () coupling the high pressure turbine to the core shaft;'}], 'a core spool comprising{'b': '45', 'a combustor () between the core spool compressor and the high pressure turbine; and'} [{'b': 160', '164', '164', '236, 'a first () of said main bearings engages a static support (; ′) and a forward hub () of the intermediate spool; and'}, {'b': '162', 'a second () of said main bearings engages the low shaft and the forward hub of the intermediate spool.'}], 'a plurality of main bearings wherein2. The engine of wherein:the forward hub extends forward from a disk of the intermediate spool compressor.3. The engine of wherein:{'b': '237', 'the ...

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Номер записи: 142
11-04-2019 дата публикации

Fan drive gear system with improved misalignment capability

Номер: US20190107004A1
Автор: William G. Sheridan
Принадлежит: United Technologies Corp

An epicyclic gear assembly includes a carrier that includes a first plate axially spaced from a second plate by a connector. A first epicyclic gear set is supported adjacent the first plate and includes a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear. A second epicyclic gear set is axially spaced from the first epicyclic gear set and supported adjacent the second plate and includes a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear. The first epicyclic gear is set and the second epicyclic gear set maintain relative intermeshing alignment during flexure-induced deformation of the carrier. The first ring gear includes a first set of teeth that extends from a first flexible flange and the second ring gear includes a second set of teeth that extends from a second flexible flange.

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Номер записи: 143
10-07-2014 дата публикации

METHOD FOR SETTING A GEAR RATIO OF A FAN DRIVE GEAR SYSTEM OF A GAS TURBINE ENGINE

Номер: US20140193238A1
Автор: Hasel Karl L., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A gas turbine engine includes a fan section including a fan that is rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0. 1. A gas turbine engine comprising:a fan section including a fan rotatable about an axis;a speed reduction device connected to the fan, wherein the speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5, wherein a bypass ratio is greater than about 11.0.2. The gas turbine engine of claim 1 , wherein the speed reduction device includes a star gear system gear ratio of at least 2.6.3. The gas turbine engine of claim 2 , wherein the speed reduction device includes a system gear ratio less than or equal to 4.1.4. The gas turbine engine of claim 3 , including the bypass ratio is less than about 22.0.5. The gas turbine engine of claim 4 , wherein the fan blade tip speed of the fan section is greater than about 1000 ft/sec and less than about 1400 ft/sec.6. The gas turbine engine of claim 1 , wherein the star system includes a sun gear claim 1 , a plurality of star gears claim 1 , a ring gear claim 1 , and a carrier.7. The gas turbine engine of claim 6 , wherein each of the plurality of star gears includes at least one bearing.8. The gas turbine engine of claim 7 , wherein the carrier is fixed from rotation.9. The gas turbine engine of claim 8 , wherein a low pressure turbine is mechanically attached to the sun gear.10. The gas turbine engine of claim 9 , wherein a fan section is mechanically attached to the ring gear.11. The gas turbine engine of claim 1 , wherein an input of the speed reduction device is rotatable in a first direction and an output of the speed reduction device is rotatable in a second direction opposite to the first direction.12. The gas turbine engine of claim 11 , including a low pressure turbine section in communication with the ...

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Номер записи: 144
13-05-2021 дата публикации

METHOD OF ASSEMBLY FOR FAN DRIVE GEAR SYSTEM WITH ROTATING CARRIER

Номер: US20210140373A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A method of assembling a fan drive gear system for a gas turbine engine according to an example of the present disclosure includes the steps of providing a unitary carrier defining a central axis and that includes spaced apart walls and circumferentially spaced mounts defining spaced apart apertures at an outer circumference of the carrier, gear pockets defined between the walls and extending to the apertures, and a central opening in at least one of the walls. The method includes the steps of inserting a plurality of intermediate gears through the central opening, moving the intermediate gears radially outwardly relative to the central axis into the gear pockets, inserting a sun gear through the central opening, moving the plurality of intermediate gears radially inwardly relative to the central axis to engage the sun gear, and coupling a fan shaft to the carrier such that the fan shaft and intermediate gears are rotatable about the central axis. A fan drive gear system is also disclosed. 1. A method of assembling a fan drive gear system for a gas turbine engine comprising the steps of:providing a unitary carrier defining a central axis and that includes spaced apart walls and circumferentially spaced mounts defining spaced apart apertures at an outer circumference of the carrier, gear pockets defined between the walls and extending to the apertures, and a central opening in at least one of the walls;inserting a plurality of intermediate gears through the central opening;moving the intermediate gears radially outwardly relative to the central axis into the gear pockets;inserting a sun gear through the central opening;moving the plurality of intermediate gears radially inwardly relative to the central axis to engage the sun gear; andcoupling a fan shaft to the carrier such that the fan shaft and intermediate gears are rotatable about the central axis.2. The method as recited in claim 1 , wherein the step of coupling the fan shaft includes attaching a torque frame to ...

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Номер записи: 145
07-05-2015 дата публикации

Turbofan Engine Main Bearing Arrangement

Номер: US20150125293A1
Автор: Frederick M. Schwarz, William G. Sheridan
Принадлежит: United Technologies Corp

A turbofan engine ( 20; 300; 400 ) comprises a fan ( 28 ), a fan drive gear system ( 60 ), a fan shaft ( 120 ) coupling the fan drive gear system to the fan, a low spool, an intermediate spool, and a core spool. The low spool engages at least three main bearings of which at least two are non-thrust bearings and at least one is a thrust bearing. The fan shaft engages at least two bearings ( 148, 150 ). The core spool engages at least two bearings ( 250, 260 ). The intermediate spool engages at least two of said bearings ( 220, 200, 230; 220, 200, 230 - 2; 200, 220, 230 - 3 ).

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Номер записи: 146
24-07-2014 дата публикации

METHOD OF ASSEMBLY FOR GAS TURBINE FAN DRIVE GEAR SYSTEM

Номер: US20140206496A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A method of assembling an epicyclic gear train comprises the steps of providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining spaced apart apertures provided at an outer circumference of the carrier. Gear pockets are provided between the walls and extend to the apertures. A central opening is in at least one of the walls. A plurality of intermediate gears are inserted through the central opening and move the intermediate gears radially outwardly into the gear pockets to extend into the apertures. A sun gear is inserted through the central opening. The plurality of intermediate gears is moved radially inwardly to engage the sun gear. 1. A method of assembling an epicyclic gear train comprising the steps of:a) providing a unitary carrier having a central axis that includes spaced apart walls and circumferentially spaced connecting structure defining spaced apart apertures provided at an outer circumference of the carrier, gear pockets provided between the walls and extending to the apertures, and a central opening in at least one of the walls;b) inserting a plurality of intermediate gears through the central opening and moving the intermediate gears radially outwardly into the gear pockets to extend into the apertures;c) inserting a sun gear through the central opening; andd) moving the plurality of intermediate gears radially inwardly to engage the sun gear.2. The method as set forth in claim 1 , wherein step d) occurs after step c).3. The method as set forth in claim 1 , wherein journal bearings are inserted within each of said intermediate gears after step d).4. The method as set forth in claim 1 , wherein a ring gear is subsequently placed on an outer periphery of the sun gears to engage the sun gears.5. The method as set forth in claim 4 , wherein said sun gear and said intermediate gears are each formed as a single gear claim 4 , and said ring gear is formed as a two-part gear. ...

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Номер записи: 147
16-04-2020 дата публикации

LIGHTWEIGHT JOURNAL SUPPORT PIN

Номер: US20200116040A1
Автор: McCune Michael E., Ogden William P., Sheridan William G.
Принадлежит:

A journal support pin to support intermediate gears for use in gas turbine engine comprises a titanium body, and an outer surface outside of the titanium body having a surface hardness that is harder than the body. A gas turbine engine and a method of forming a journal support pin to support intermediate gears for use in gas turbine engine are also disclosed. 1. A journal support pin to support intermediate gears for use in gas turbine engine comprising:a titanium body, and an outer surface outside of said titanium body having a surface hardness that is harder than said body;said outer surface is provided by one of a steel sleeve, a coating, nitriding or high velocity oxyfuel spray; andoil supply holes extend from a central bore in said body through said outer surface.2. The journal support pin as set forth in claim 1 , wherein said oil supply holes extend through said outer surface at a recess claim 1 , wherein a thickness of said outer surface is thinner than at axial ends of said titanium body.3. The journal support pin as set forth in claim 2 , wherein said recess extends only over a limited circumferential portion of said outer surface.4. The journal support pin as set forth in claim 1 , wherein said outer surface is provided by a steel sleeve.5. The journal support pin as set forth in claim 4 , wherein said steel sleeve is pinned to said body to secure said sleeve to said body.6. The journal support pin as set forth in claim 1 , wherein said outer surface is provided by a coating.7. The journal support pin as set forth in claim 6 , wherein said coating is one of silver claim 6 , steel and titanium nitride.8. The journal support pin as set forth in claim 1 , wherein said outer surface is provided by nitriding.9. The journal support pin as set forth in claim 1 , wherein said outer surface is provided by high velocity oxyfuel spray.10. A gas turbine engine comprising:a fan and a fan drive turbine, said fan drive turbine driving said fan through a gear reduction; ...

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Номер записи: 148
10-05-2018 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20180128185A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan shaft arranged along an engine central axis, a frame supporting the fan shaft, a gear system rotatably coupled with the fan shaft, and a flexible coupling at least partially supporting the gear system. The flexible coupling defines, with respect to the engine central axis, a torsional stiffness TS and a lateral stiffness LS such that a ratio of TS/LS is greater than or equal to about 2.0 to reduce loads on the gear system from misalignment of the gear system with respect to the engine central axis. 1. A gas turbine engine comprising:a fan shaft arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a frame lateral stiffness;a gear system rotatably coupled to the fan shaft, the gear system having a gear reduction ratio that is greater than 2.3; anda first, non-rotatable flexible coupling and a second, rotatable flexible coupling supporting the gear system, the first flexible coupling and the second flexible coupling being subject to, with respect to the engine central axis, parallel offset guided end motion,the first flexible coupling having a Stiffness D that is torsional stiffness under the parallel offset guided end motion, and the first flexible coupling has a ratio of frame lateral stiffness to Stiffness D that is in a range of 0.25 to 0.50.2. The gas turbine engine as recited in claim 1 , wherein the gear system includes a sun gear in meshed engagement with multiple intermediate gears that are rotatably mounted on bearings in a non-rotatable carrier claim 1 , each intermediate gear is in meshed engagement with a rotatable ring gear claim 1 , the sun gear is rotatably coupled to the fan shaft claim 1 , and the first claim 1 , non-rotatable flexible coupling is coupled with the non-rotatable carrier.3. The gas turbine engine as recited in claim 2 , wherein the gear system is coupled through an input shaft to a low pressure turbine claim 2 , the low pressure turbine having a pressure ratio of ...

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Номер записи: 149
10-05-2018 дата публикации

FLEXIBLE COUPLING FOR GEARED TURBINE ENGINE

Номер: US20180128186A1
Автор: Otto John R., Sheridan William G.
Принадлежит:

A gas turbine engine includes a fan shaft arranged along an engine central axis, a frame supporting the fan shaft, a gear system rotatably coupled with the fan shaft, and a flexible coupling at least partially supporting the gear system. The flexible coupling defines, with respect to the engine central axis, a torsional stiffness TS and a lateral stiffness LS such that a ratio of TS/LS is greater than or equal to about 2.0 to reduce loads on the gear system from misalignment of the gear system with respect to the engine central axis. 1. A gas turbine engine comprising:a fan shaft arranged along an engine central axis;a frame supporting the fan shaft, the frame defining a frame lateral stiffness;a gear system rotatably coupled to the fan shaft, the gear system having a gear reduction ratio that is greater than 2.3; anda first, non-rotatable flexible coupling and a second, rotatable flexible coupling supporting the gear system, the first flexible coupling and the second flexible coupling being subject to, with respect to the engine central axis, parallel offset guided end motion,the second flexible coupling having a Stiffness D that is torsional stiffness under the parallel offset guided end motion, and the second flexible coupling has a ratio of frame lateral stiffness to Stiffness D that is in a range of 2 to 100.2. The gas turbine engine as recited in claim 1 , wherein the gear system includes a sun gear in meshed engagement with multiple intermediate gears that are rotatably mounted on bearings in a non-rotatable carrier claim 1 , each intermediate gear is in meshed engagement with a rotatable ring gear claim 1 , the sun gear is rotatably coupled to the fan shaft claim 1 , and the first claim 1 , non-rotatable flexible coupling is coupled with the non-rotatable carrier.3. The gas turbine engine as recited in claim 2 , wherein the gear system is coupled through an input shaft to a low pressure turbine claim 2 , the low pressure turbine having a pressure ratio of ...

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Номер записи: 150
02-05-2019 дата публикации

Turbofan Engine Bearing and Gearbox Arrangement

Номер: US20190128182A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A turbofan engine comprising a fan shaft configured to rotate about an axis of the engine. A fan drive gear system is configured to drive the fan shaft. A first spool comprises a high pressure turbine and a high pressure compressor. A second spool comprises a lower pressure turbine, a lower pressure compressor, and a shaft coupling the lower pressure turbine to the intermediate pressure compressor. The engine has a plurality of main bearings. The lower pressure compressor has a plurality of disks and a forward hub mounted to a forwardmost disk. 2. The engine of wherein:{'b': 25', '336, 'of the main bearings, at least one is a shaft-engaging bearing engaging a driving shaft (; ) coupled to the fan drive gear system;'}{'b': 160', '340', '640', '550, 'sub': 'B', 'a closest (; ; ) of the shaft-engaging bearings engaging the driving shaft behind the fan drive gear system has a centerplane () and a characteristic radius (R);'}{'b': 530', '540, 'sub': 'B', 'the half angle (θ) of a virtual cone () intersecting the core flowpath inboard boundary at the gear system centerplane () and said closest of the shaft-engaging bearings at the characteristic radius (R) is greater than 32°; and'}{'sub': R', 'R, 'a hub-to-tip ratio (H:F) of the fan is less than 0.38.'}3202. The engine of wherein a universal joint () couples the driving shaft to the fan drive gear system.4. The engine of wherein the angle (θ) is 33° to 68°.5. The engine of wherein the hub-to-tip ratio (H:F) is 0.24 to 0.33.6. The engine of wherein the angle (θ) is greater than 40°.7. The engine of wherein the angle (θ) is greater than 50°.825. The engine of wherein the closest of the shaft-engaging bearings behind the fan drive gear system is a low spool bearing directly coupling the shaft () to a case immediately behind the gear system.9148150. The engine of wherein at least one of said main bearings is a fan bearing ( claim 1 , ) engaging the fan shaft forward of the gear system centerplane.10. The engine of wherein the ...

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Номер записи: 151
19-05-2016 дата публикации

Gear carrier flex mount lubrication

Номер: US20160138422A1
Автор: James B. Coffin, Michael E. McCune, William G. Sheridan
Принадлежит: United Technologies Corp

An exemplary method of lubricating a turbomachine interface includes, among other things, securing a carrier relative to a torque frame using a flexure pin, and lubricating an interface of the flexure pin using a lubricant that has moved through a lubricant passage in the carrier and a lubricant passage in the flexure pin.

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Номер записи: 152
19-05-2016 дата публикации

TURBOFAN ENGINE FRONT SECTION

Номер: US20160138425A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит:

A front section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan section including a fan hub. The fan hub includes a hub diameter supporting a plurality of fan blades including a tip diameter. A transitional entrance passage is configured to communicate flow between the fan section and a compressor section. The transitional entrance passage includes an inlet disposed at an inlet diameter and an outlet to the compressor section disposed at an outlet diameter. A method of designing a gas turbine engine is also disclosed. 1. A front section for a gas turbine engine comprising:a fan section including a fan hub, the fan hub including a hub diameter supporting a plurality of fan blades including a tip diameter, a ratio of the hub diameter to the tip diameter being between about 0.24 and about 0.36; anda transitional entrance passage configured to communicate flow between the fan section and a compressor section, the transitional entrance passage including an inlet disposed at an inlet diameter and an outlet to the compressor section disposed at an outlet diameter, a ratio of the hub diameter to the inlet diameter being between about 0.65 and about 0.95.2. The front section as recited in claim 1 , wherein the plurality of fan blades is less than twenty (20) fan blades.3. The front section as recited in claim 2 , wherein the ratio of the hub diameter to the inlet diameter is between about 0.70 and about 0.90.4. The front section as recited in claim 1 , wherein the fan section is configured to deliver air into a bypass duct claim 1 , and a portion of air into the compressor section claim 1 , with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section claim 1 , and the bypass ratio being greater than about 10.5. The front section as recited in claim 4 , wherein a pressure ratio across the fan section is less than about 1.5.6. ...

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Номер записи: 153
28-05-2015 дата публикации

Geared Turbofan Engine Gearbox Arrangement

Номер: US20150143794A1
Автор: Schwarz Frederick M., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A three-spool turbofan engine () has a variable fan nozzle (). The fan blades have a peak tip radius Rand an inboard leading edge radius Rat an inboard boundary of the flowpath. A ratio of Rto Ris less than about 0.40. 120. A turbofan engine () comprising:{'b': '28', 'a fan () having a plurality of blades;'}{'b': '60', 'a transmission() configured to drive the fan; and'}three spools, wherein:{'sub': 'T', 'the fan blades have a peak tip radius R;'}{'sub': 'H', 'the fan blades have an inboard leading edge radius Rat an inboard boundary of the flowpath;'}{'sub': H', 'T, 'a ratio of Rto Ris less than about 0.40;'} [ [{'b': '50', 'a first pressure turbine (); and'}, {'b': '56', 'a first shaft () coupling the first pressure turbine to the transmission;'}], 'a first spool comprising, [{'b': '48', 'a second pressure turbine ();'}, {'b': '42', 'a first compressor (); and'}, {'b': '54', 'a second spool shaft () coupling the second pressure turbine to the second spool compressor;'}], 'a second spool comprising;'}, [{'b': '46', 'a third pressure turbine ();'}, {'b': '44', 'a second compressor (); and'}, {'b': '52', 'a core shaft () coupling the third pressure turbine to second compressor; and'}], 'a core spool comprising], 'said three spools comprise{'b': '45', 'a combustor () is between the second compressor and the third pressure turbine;'} [{'b': 160', '164', '164', '236, 'a first () of said main bearings engages a static support (; ′) and a forward hub () of the second spool; and'}, {'b': '162', 'a second () of said main bearings engages the first shaft and the forward hub of the second spool;'}], 'the engine comprises a plurality of main bearings whereinthe fan is a single-stage fan; and{'b': '82', 'a ring gear () of the transmission is mounted to rotate with the fan as a unit.'}2. The engine of wherein:each fan blade has a leading edge and a trailing edge;{'b': 159', '506', '510, 'a splitter () is positioned along a flowpath through the engine and having a leading rim ...

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Номер записи: 154
17-05-2018 дата публикации

METHOD OF ASSEMBLY FOR GAS TURBINE FAN DRIVE GEAR SYSTEM

Номер: US20180135523A1
Автор: McCune Michael E., Sheridan William G.
Принадлежит:

A method of assembling an epicyclic gear train according to an example of the present disclosure includes, among other things, the steps of providing a carrier having a central axis that includes spaced apart side walls and circumferentially spaced connecting structure defining mounts that interconnect the side walls, spaced apart apertures provided at an outer circumference of the carrier, gear pockets provided between the side walls and extending to the apertures, and a central opening in at least one of the walls, providing oil baffles between the side walls, the oil baffles including ends that abut the mounts, inserting a plurality of intermediate gears through the central opening, and then moving the intermediate gears radially outwardly into the gear pockets to extend into the apertures, inserting a sun gear through the central opening, and moving the plurality of intermediate gears radially inwardly to engage the sun gear. 1. A method of assembling an epicyclic gear train comprising the steps of:providing a carrier having a central axis that includes spaced apart side walls and circumferentially spaced connecting structure defining mounts that interconnect the side walls, spaced apart apertures provided at an outer circumference of the carrier, gear pockets provided between the side walls and extending to the apertures, and a central opening in at least one of the walls;providing oil baffles between the side walls, the oil baffles including ends that abut the mounts;inserting a plurality of intermediate gears through the central opening, and then moving the intermediate gears radially outwardly into the gear pockets to extend into the apertures;inserting a sun gear through the central opening; andmoving the plurality of intermediate gears radially inwardly to engage the sun gear.2. The method as set forth in claim 1 , wherein the step of inserting the sun gear occurs after the step of providing the oil baffles.3. The method as set forth in claim 2 , wherein ...

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Номер записи: 155
17-05-2018 дата публикации

GEAR SYSTEM ARCHITECTURE FOR GAS TURBINE ENGINE

Номер: US20180135524A1
Автор: Sheridan William G.
Принадлежит:

A disclosed gas turbine engine includes a fan including a plurality of fan blades rotatable about an axis, a bypass duct, a compressor section and a bypass ratio greater than 10. A combustor in fluid communication with the compressor section and a fan drive turbine in communication with the combustor. A gear system provides a speed reduction between the fan drive turbine and the fan and transfer power input from the fan drive turbine to the fan at an efficiency greater than 98%. 1. A gas turbine engine comprising:a fan including less than twenty (20) fan blades rotatable about an axis;a compressor section;a combustor in communication with the compressor section;a bypass duct and a bypass ratio of a portion of air delivered into the bypass duct divided by the amount of air delivered into the compressor section is greater than ten (10);a fan drive turbine in communication with the combustor, the fan drive turbine comprising less than six (6) fan drive turbine rotors, an inlet having an inlet pressure, an outlet prior to any exhaust nozzle and having an outlet pressure, a pressure ratio of the inlet pressure to the outlet pressure greater than about five (5), and a ratio of between the number of fan blades and the number of fan drive turbine rotors is between 3.3 and 8.6;a bearing system supporting rotation of an inner shaft driven by the fan drive turbine and an outer shaft driven by a second turbine;a fan drive gear system providing a speed reduction between the fan drive turbine and the fan of greater than 2.3, the fan drive gear system transferring power input from the fan drive turbine to the fan at an efficiency greater than 98% and less than 100%;a lubrication system circulating lubricant to at least the bearing system and the fan drive gear system, the lubrication system has a maximum capacity for removing thermal energy from the lubricant equal of no more than 2% of power input into the fan drive gear system by the fan drive turbine during operation of the ...

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Номер записи: 156
30-04-2020 дата публикации

Wet-Face/Dry-Face Seal and Methods of Operation

Номер: US20200132196A1
Автор: Davis Todd A., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A seal system has: a first member; a seal carried by the first member and having a seal face; and a second member rotatable relative to the first member about an axis. The second member has: a seat, the seat having a seat face in sliding sealing engagement with the seal face; and a circumferential array of passageway legs open to the seat face; and an oil pump for delivering oil via one or more first outlets to the passageway legs in at least a first mode of operation. The oil pump is coupled to one or more second outlets to deliver oil to a backside of the seat in at least a second mode of operation. 2. The seal system of wherein:the seal is a carbon seal.3. The seal system of wherein:the seat is steel.4. The seal system of wherein:the seal and seat are full annular.5. The seal system of further comprising:a spring biasing the seal into engagement with the seat.6. The seal system of further comprising:one or more valves for switching between the first mode and the second mode.7. The seal system of wherein:the one or more valves comprises a first valve positioned along a flowpath from the pump to the one or more first outlets downstream of a branching to the one or more second outlets.8. The seal system of wherein:the one or more valves consists of a single valve;the one or more first outlets consists of a single first outlet; andthe one or more second outlets consists of a single second outlet.9. The seal system of wherein:the one or more valves are positioned to allow flow through the second outlets in the first mode and the second mode and block flow through the first outlets in the second mode.10. The seal system of further comprising a nozzle body wherein:one of the first outlets is in a first insert in the nozzle body; andone of the second outlets is in a second insert in the nozzle body.11. The seal system of further comprising a valve in the nozzle body positioned to selectively block flow through the first insert.12. A gas turbine engine including the seal ...

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Номер записи: 157
14-08-2014 дата публикации

COMPOUND STAR GEAR SYSTEM WITH ROLLING ELEMENT BEARINGS

Номер: US20140227084A1
Автор: Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

A compound star gear system includes a sun gear rotatable about a first axis that drives a first plurality of star gears rotatable about a plurality of fixed axes. The first plurality of star gears drives a second plurality of star gears spaced axially apart from the first plurality of star gears. The second plurality of star gears drive the ring gear that in turn drives a fan drive shaft. 1. A fan drive gear assembly comprising:a sun gear rotatable about a first axis;a first plurality of star gears driven by the sun gear and rotatable about a plurality of fixed axes;a second plurality of star gears rotatable with the first plurality of star gears about the plurality of fixed axes;a ring gear driven by the second plurality of star gears; anda fan drive shaft driven by the ring gear.2. The fan drive gear assembly as recited in claim 1 , wherein sun gear and the first plurality of star gears are engaged within a first plane and the second plurality of star gears and the ring ear are engaged within a second plane spaced axially apart from the first plane.3. The fan drive gear assembly as recited in claim 1 , wherein each of the first plurality of star gears comprises a first diameter and each of the second plurality of star gears comprises a second diameter greater than the first diameter.4. The fan drive gear assembly as recited in claim 1 , wherein the first plurality of star gears and the second plurality of star gears are supported for rotation by corresponding pluralities of rolling elements.5. The fan drive gear assembly as recited in claim 4 , wherein the rolling elements comprise cylindrical roller bearings.6. The fan drive gear assembly as recited in claim 4 , wherein the rolling elements comprise spherical roller bearings.7. The fan drive gear assembly as recited in claim 4 , wherein the rolling elements comprise tapered roller bearings.8. The fan drive gear assembly as recited in claim 4 , including a plurality of shaft supporting rotation of the first and ...

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Номер записи: 158
21-08-2014 дата публикации

PLANETARY GEAR SYSTEM ARRANGEMENT WITH AUXILIARY OIL SYSTEM

Номер: US20140230452A1
Автор: McCune Michael E., Portlock Lawrence E., Sheridan William G.
Принадлежит: UNITED TECHNOLOGIES CORPORATION

In one exemplary embodiment, a gas turbine engine includes a fan, a speed reduction device driving the fan, and a lubrication system for lubricating components across a rotation gap. The lubrication system includes a lubricant input. A stationary first bearing receives lubricant from the lubricant input and has a first race in which lubricant flows. A second bearing for rotation is within the first bearing. The second bearing has a first opening in registration with said first race such that lubricant may flow from the first race through the first opening into a first conduit. The first bearing also has a second race into which lubricant flows. The second bearing has a second opening in registration with the second race such that lubricant may flow from the second race through the second opening into a second conduit. The first and second conduits deliver lubricant to distinct locations. 1. A gas turbine engine comprising:a fan;a speed reduction device driving said fan;a lubrication system for lubricating components across a rotation gap, the lubrication system including a lubricant input, a stationary first bearing receiving lubricant from said lubricant input and having a first race in which lubricant flows, and a second bearing for rotation within said first bearing, said second bearing having a first opening in registration with said first race such that lubricant may flow from said first race through said first opening into a first conduit; andsaid first bearing also having a second race into which lubricant flows, and said second bearing having a second opening in registration with said second race such that lubricant may flow from said second race through said second opening into a second conduit, with said first and second conduits delivering lubricant to distinct locations.2. The gas turbine engine of claim 1 , further comprising a first compressor rotor and a second compressor rotor.3. The gas turbine engine of claim 2 , further comprising a first turbine ...

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Номер записи: 159