WO2015047489A1 - Configuration de palier d'arbre de moteur à turbine à gaz - Google Patents

Configuration de palier d'arbre de moteur à turbine à gaz Download PDF

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Publication number
WO2015047489A1
WO2015047489A1 PCT/US2014/043195 US2014043195W WO2015047489A1 WO 2015047489 A1 WO2015047489 A1 WO 2015047489A1 US 2014043195 W US2014043195 W US 2014043195W WO 2015047489 A1 WO2015047489 A1 WO 2015047489A1
Authority
WO
WIPO (PCT)
Prior art keywords
gas turbine
turbine engine
inlet
engine according
compressor
Prior art date
Application number
PCT/US2014/043195
Other languages
English (en)
Inventor
Brian D. Merry
Gabriel L. Suciu
Karl L. Hasel
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from US14/012,773 external-priority patent/US8863491B2/en
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to EP14849357.0A priority Critical patent/EP3027864A4/fr
Publication of WO2015047489A1 publication Critical patent/WO2015047489A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
    • F01D25/162—Bearing supports
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/06—Arrangements of bearings; Lubricating
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
    • F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type

Definitions

  • Turbomachines such as gas turbine engines, typically include a fan section, a turbine section, a compressor section, and a combustor section.
  • the fan section drives air along a core flow path into the compressor section.
  • the compressed air is mixed with fuel and combusted in the combustor section.
  • the products of combustion are expanded in the turbine section.
  • a typical jet engine has two or three spools, or shafts, that transmit torque between the turbine and compressor sections of the engine.
  • Each of these spools is typically supported by two bearings.
  • One bearing for example, a ball bearing, is arranged at a forward end of the spool and is configured to react to both axial and radial loads.
  • Another bearing for example, a roller bearing is arranged at the aft end of the spool and is configured to react only to radial loads. This bearing arrangement fully constrains the shaft except for rotation, and axial movement of one free end is permitted to accommodate engine axial growth.
  • a core inlet typically controls flow of air into the core flow path.
  • the flow of air moves from the core inlet to a compressor section inlet.
  • the relative radial positions of the core inlet and the compressor section inlet may influence flow through the core and a profile of the turbomachine.
  • a gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path.
  • a geared architecture is arranged within the inlet case.
  • a shaft provides a rotational axis.
  • a hub is operatively supported by the shaft.
  • a rotor is connected to the hub and supports a compressor section. The compressor section is arranged axially between the inlet case flow path and the intermediate case flow path.
  • a bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.
  • a core inlet includes the inlet case flow path and has a radially inner boundary that is spaced a first radial distance from the rotational axis.
  • a compressor section inlet has a radially inner boundary that is spaced a second radial distance from the rotational axis.
  • a ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.
  • the radially inner boundary of the core inlet is at a location of a core inlet stator.
  • the radially inner boundary of the compressor section inlet is at a location of a compressor rotor.
  • the compressor rotor is a first stage rotor of a low-pressure compressor.
  • the core inlet is an inlet to the core housing.
  • an inlet flow of the compressor section is configured to be from about 30 lb/sec/ft 2 to about 37 lb/sec/ft 2 when the engine is operating at a cruise speed.
  • a turbine inlet temperature of a high-pressure turbine within the engine is configured to be from about 2,000°F to about 2,600°F when the engine is operating at a cruise speed.
  • a tip speed of a blade array in the compressor section during engine operation is configured to be from about 1,050 fps to about 1,350 fps.
  • a fan section is driven by the geared architecture that is driven by the shaft that rotates a compressor rotor within the compressor section.
  • the geared architecture has a gear reduction ratio of greater than about 2.3.
  • the inlet case includes a first inlet case portion that defines the inlet case flow path.
  • a bearing support portion is removably secured to the inlet case portion.
  • the bearing is mounted to the bearing support portion.
  • the intermediate case includes an intermediate case portion that defines the intermediate case flow path.
  • a bearing support portion is removably secured to the intermediate case portion.
  • the bearing is mounted to the bearing support portion.
  • the bearing is a ball bearing.
  • the bearing is a first bearing and further comprising a second bearing that supports the shaft relative to the other of the intermediate case and the inlet case.
  • first and second bearings are arranged in separate sealed lubrication compartments.
  • the geared architecture is coupled to the shaft.
  • a fan is coupled to and rotationally driven by the geared architecture.
  • the shaft includes a main shaft and a flex shaft.
  • the flex shaft is secured to the main shaft at a first end and including a second end opposite the first end.
  • the geared architecture includes a sun gear supported on the second end.
  • the shaft includes a hub that is secured to the main shaft.
  • the compressor section includes a rotor mounted to the hub.
  • the geared architecture includes a torque frame that supports multiple circumferentially arranged star gears that intermesh with the sun gear. The torque frame is secured to the inlet case.
  • the rotor supports multiple compressor stages.
  • the bearing is axially aligned with and radially inward of one of the compressor stages.
  • the compressor section includes a variable stator vane array.
  • the geared architecture is arranged in the lubrication compartment.
  • the core housing includes a core inlet stator that has a stator root that is spaced a first radial distance from the rotational axis.
  • the compressor section includes a compressor blade having a blade root that is spaced a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.
  • stator root is radially aligned with a radially inner boundary of a core flow path through the gas turbine engine.
  • the blade root is radially aligned with a radially inner boundary of a core flow path through the gas turbine engine.
  • the core inlet stator is positioned within an inlet to a core section of the gas turbine engine.
  • Figure 1 schematically illustrates an embodiment of a gas turbine engine.
  • Figure 2 is a cross-sectional view of a front architecture of the gas turbine engine embodiment shown in Figure 1.
  • Figure 3 shows a close-up view of a core inlet portion of the Figure 1 gas turbine engine embodiment.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath B while the compressor section 24 drives air along a core flowpath C (as shown in Figure 2) for compression and communication into the combustor section 26 then expansion through the turbine section 28.
  • the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid-turbine frame 57 supports one or more bearing systems 38 in the turbine section 28.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A, which is collinear with their longitudinal axes.
  • the core airflow C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
  • the engine 20 in one example a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10).
  • the example speed reduction device is a geared architecture 48 however other speed reducing devices such as fluid or electromechanical devices are also within the contemplation of this disclosure.
  • the example geared architecture 48 is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3, or more specifically, a ratio of from about 2.2 to about 4.0.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5:1.
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • a significant amount of thrust is provided by the bypass flow B due to the high bypass ratio.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet.
  • TSFC Thrust Specific Fuel Consumption
  • TSFC is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
  • Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(T a m b i e n t °R) / 518.7 °R) ° '5 ].
  • the "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second.
  • a core housing 60 includes an inlet case 62 and an intermediate case 64 that respectively provide an inlet case flowpath 63 and a compressor case flowpath 65.
  • the core housing may include additional cases.
  • the compressor section as a whole may include any number of cases.
  • the inlet and compressor case flowpaths 63, 65 in part, define a core flowpath through the engine 20, which directs a core flow C.
  • the intermediate case 64 includes multiple components, including the intermediate case portion 66, and the bearing support 68 in the example, which are removably secured to one another.
  • the bearing support portion 68 has a first bearing 70 mounted thereto, which supports the inner shaft 40 for rotation relative to the intermediate case 64.
  • the first bearing 70 is a ball bearing that constrains the inner shaft 40 against axial and radial movement at a forward portion of the inner shaft 40.
  • the first bearing 70 is arranged within a bearing compartment 71.
  • the inner shaft 40 is constructed of multiple components that include, for example, a main shaft 72, a hub 74 and a flex shaft 76, which are clamped together by a nut 80 in the example.
  • the first bearing 70 is mounted on the hub 74 (i.e., low pressure compressor hub).
  • the flex shaft 76 includes first and second opposing ends 82, 84. The first end 82 is splined to the hub 74, and the second end 84 is splined to and supports a sun gear 86 of the geared architecture 48. Bellows 78 in the flex shaft 76 accommodate vibration in the geared architecture 48.
  • the geared architecture includes star gears 88 arranged circumferentially about and intermeshing with the sun gear 86.
  • a ring gear 90 is arranged circumferentially about and intermeshes with the star gears 88.
  • a fan shaft 92 is connected to the ring gear 90 and the fan 42 ( Figure 1).
  • a torque frame 94 supports the star gears 88 and grounds the star gears 88 to the housing 60. In operation, the inner shaft 40 rotationally drives the fan shaft 92 with the rotating ring gear 90 through the grounded star gears 88.
  • the low pressure compressor 44 includes multiple compressor stages arranged between the inlet and intermediate case flowpaths 63, 65, for example, first and second compressor stages 98, 100, that are secured to the hub 74 by a rotor 96.
  • the first bearing 70 is axially aligned with one of the first and second compressor stages 98, 100.
  • a variable stator vane array 102 is arranged upstream from the first and second compressor stages 98, 100.
  • Struts 104 are arranged upstream from the variable stator vane array 102.
  • An array of fixed stator vanes 106 may be provided axially between the first and second compressor stages 98, 100.
  • the inlet case 62 includes inlet case portions 108, and bearing support 110, which are removably secured to one another.
  • the bearing support portion 110 and torque frame 94 are secured to the inlet case portion 108 at a joint 109.
  • the bearing support portion 110 supports a second bearing 112, which is a rolling bearing in one example.
  • the second bearing 112 is retained on the hub 74 by a nut 113, for example, and is arranged radially outward from the flex shaft 76 and radially between the torque frame 94 and flex shaft 76.
  • the second bearing 112 is axially aligned with and radially inward of the variable stator vane array 102.
  • the geared architecture 48 and the second bearing 112 are arranged in a lubrication compartment 114, which is separate from the bearing compartment 71 in the example.
  • the core flow path of the example engine 20 begins at a core inlet 160 and extends through and past the low-pressure compressor 44.
  • the core inlet 160 has a radially inner boundary 162 and a radially outer boundary 166.
  • a core inlet stator 170 is located at or near the core inlet 160.
  • the core inlet stator 170 attaches to a core case 174 at the radially inner boundary 162.
  • the core inlet stator 170 attaches to an inlet case 178 at the radially outer boundary 166.
  • the core inlet stator 170 extends radially across the core flow path C.
  • the radially inner boundary 162 is positioned a radial distance Di from the axis A.
  • the distance Di in this example, also corresponds to the radial distance between a root 164 of the core inlet stator 170 and the axis A.
  • the root 164 of the core inlet stator 170 is radially aligned with the radially inner boundary 162 of the core flow path C.
  • the compressor section inlet 182 is an inlet to the low-pressure compressor 44 of the compressor section 24.
  • the compressor inlet 182 extends from a radially inner boundary 186 to a radially outer boundary 190.
  • a blade 198 of a rotor within the low-pressure compressor 44 extends from a root 202 to a tip 206.
  • the blade 198 is located at or near the compressor inlet 182.
  • the blade 198 part of a compressor rotor within a first stage of the compressor section 24.
  • the blade 198 is thus part of a first stage rotor, or a leading blade of the compressor section 24 relative to a direction of flow along the core flow path C.
  • the blade 198 represents the axial position where air enters the compressor section 24 of the core flow path C.
  • the blade 198 extends radially across the core flow path C.
  • the radially inner boundary 186 is positioned a radial distance D 2 from the axis A.
  • the distance D 2 in this example, also corresponds to the radial distance between the root 202 of the blade 198 and the axis A.
  • the root 202 is radially aligned with the radially inner boundary 186 of the core flow path C.
  • a preferred ratio range of the distance D 2 to the distance Di spans from about 0.65 to about 0.9, which provides a relatively low profile core flow path contour.
  • High profile flow path contours have greater differences between D 2 and
  • High profile flow path contours introduce discontinuities that undesirably disrupt the airflow and undesirably add weight to the engine 20.
  • the ratio range of about 0.65 to about 0.9 is made possible, in part, by the incorporation of the geared architecture 48 into the engine 20.
  • the "hump" in this example is generally area 200.
  • the distance D 2 may be larger as compared to bearing arrangements which have bearings axially offset from the compressor section.
  • the axially compact low pressure compressor and bearing arrangement also minimizes discontinuities in the flow path contour while reducing the axial length of the engine.
  • Other characteristics of the engine having this ratio may include the engine 20 having a specific inlet flow of the low pressure compressor at cruising speeds to be between about 30 lb/sec/ft 2 and about 37 lb/sec/ft 2 .
  • the specific inlet flow is the amount of flow moving into the compressor section 24 and specifically, in this example, into a compressor inlet 182 and through the compressor section 24.
  • Another characteristic of the example engine 20 is that the cruise speeds of the example engine are generally Mach numbers of about 0.7 to about 0.9.
  • a temperature at an inlet to the high-pressure turbine 54 may be from about 2,000°F (1093.33°C) to about 2,600°F (1426.66°C). Maintaining temperatures within this range balances good fuel consumption, low engine weight, and low engine maintenance costs.
  • a tip speed of blades in a rotor of the low-pressure compressor 44 may be from about 1,050 fps (320 m/s) to about 1,350 fps (411 m/s).
  • the geared architecture 48 of the engine 20 may have a gear ratio of greater than about 2.3, or more specifically, a ratio of from about 2.2 to about 4.0.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne un moteur à turbine à gaz comprenant un logement central qui comprend un carter d'admission et un carter intermédiaire qui fournissent respectivement un trajet d'écoulement de carter d'admission et un trajet d'écoulement de carter intermédiaire. Une architecture à engrenages est agencée dans le carter d'admission. Un arbre fournit un axe de rotation. Un moyeu est fonctionnellement supporté par l'arbre. Un rotor est raccordé au moyeu et supporte une section de compresseur. La section de compresseur est agencée de façon axiale entre le trajet d'écoulement de carter d'admission et le trajet d'écoulement de carter intermédiaire. Un palier est monté sur le moyeu et supporte l'arbre par rapport au carter intermédiaire ou au carter d'admission.
PCT/US2014/043195 2013-07-31 2014-06-19 Configuration de palier d'arbre de moteur à turbine à gaz WO2015047489A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14849357.0A EP3027864A4 (fr) 2013-07-31 2014-06-19 Configuration de palier d'arbre de moteur à turbine à gaz

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
US201361860337P 2013-07-31 2013-07-31
US61/860,337 2013-07-31
US14/012,773 US8863491B2 (en) 2012-01-31 2013-08-28 Gas turbine engine shaft bearing configuration
US14/012,773 2013-08-28

Publications (1)

Publication Number Publication Date
WO2015047489A1 true WO2015047489A1 (fr) 2015-04-02

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ID=52744312

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2014/043195 WO2015047489A1 (fr) 2013-07-31 2014-06-19 Configuration de palier d'arbre de moteur à turbine à gaz

Country Status (2)

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EP (1) EP3027864A4 (fr)
WO (1) WO2015047489A1 (fr)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107061008A (zh) * 2015-11-17 2017-08-18 通用电气公司 燃气涡轮发动机风扇
US11225975B2 (en) 2015-11-17 2022-01-18 General Electric Company Gas turbine engine fan
US11401831B2 (en) 2012-01-31 2022-08-02 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
US11486269B2 (en) * 2012-01-31 2022-11-01 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
US11566586B2 (en) 2012-01-31 2023-01-31 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
US11674435B2 (en) 2021-06-29 2023-06-13 General Electric Company Levered counterweight feathering system
US11795964B2 (en) 2021-07-16 2023-10-24 General Electric Company Levered counterweight feathering system

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US4003199A (en) * 1976-03-01 1977-01-18 General Motors Corporation Turbine engine with air brake
US4827712A (en) * 1986-12-23 1989-05-09 Rolls-Royce Plc Turbofan gas turbine engine
US7493753B2 (en) * 2005-10-19 2009-02-24 General Electric Company Gas turbine engine assembly and methods of assembling same
US7694505B2 (en) * 2006-07-31 2010-04-13 General Electric Company Gas turbine engine assembly and method of assembling same

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US9957918B2 (en) * 2007-08-28 2018-05-01 United Technologies Corporation Gas turbine engine front architecture
US8511987B2 (en) * 2009-11-20 2013-08-20 United Technologies Corporation Engine bearing support
US8402741B1 (en) * 2012-01-31 2013-03-26 United Technologies Corporation Gas turbine engine shaft bearing configuration

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US4003199A (en) * 1976-03-01 1977-01-18 General Motors Corporation Turbine engine with air brake
US4827712A (en) * 1986-12-23 1989-05-09 Rolls-Royce Plc Turbofan gas turbine engine
US7493753B2 (en) * 2005-10-19 2009-02-24 General Electric Company Gas turbine engine assembly and methods of assembling same
US7694505B2 (en) * 2006-07-31 2010-04-13 General Electric Company Gas turbine engine assembly and method of assembling same

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Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11401831B2 (en) 2012-01-31 2022-08-02 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
US11486269B2 (en) * 2012-01-31 2022-11-01 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
US11566586B2 (en) 2012-01-31 2023-01-31 Raytheon Technologies Corporation Gas turbine engine shaft bearing configuration
CN107061008A (zh) * 2015-11-17 2017-08-18 通用电气公司 燃气涡轮发动机风扇
EP3208448A1 (fr) * 2015-11-17 2017-08-23 General Electric Company Ventilateur de moteur à turbine à gaz
US10371096B2 (en) 2015-11-17 2019-08-06 General Electric Company Gas turbine engine fan
CN107061008B (zh) * 2015-11-17 2019-12-13 通用电气公司 燃气涡轮发动机
CN111120101A (zh) * 2015-11-17 2020-05-08 通用电气公司 燃气涡轮发动机
US11225975B2 (en) 2015-11-17 2022-01-18 General Electric Company Gas turbine engine fan
CN111120101B (zh) * 2015-11-17 2024-01-16 通用电气公司 燃气涡轮发动机
US11674435B2 (en) 2021-06-29 2023-06-13 General Electric Company Levered counterweight feathering system
US11795964B2 (en) 2021-07-16 2023-10-24 General Electric Company Levered counterweight feathering system

Also Published As

Publication number Publication date
EP3027864A4 (fr) 2017-03-22
EP3027864A1 (fr) 2016-06-08

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