WO2015017042A1 - Forme de trajet d'écoulement de compresseur basse pression (lpc) avec configuration de palier de moteur de turbine à gaz - Google Patents

Forme de trajet d'écoulement de compresseur basse pression (lpc) avec configuration de palier de moteur de turbine à gaz Download PDF

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Publication number
WO2015017042A1
WO2015017042A1 PCT/US2014/043184 US2014043184W WO2015017042A1 WO 2015017042 A1 WO2015017042 A1 WO 2015017042A1 US 2014043184 W US2014043184 W US 2014043184W WO 2015017042 A1 WO2015017042 A1 WO 2015017042A1
Authority
WO
WIPO (PCT)
Prior art keywords
gas turbine
turbine engine
engine according
low pressure
shaft
Prior art date
Application number
PCT/US2014/043184
Other languages
English (en)
Inventor
Brian D. Merry
Gabriel L. Suciu
Lisa I. Brilliant
Becky E. Rose
Yuan Dong
Stanley J. Balamucki
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 US14/067,354 external-priority patent/US9038366B2/en
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to EP14831790.2A priority Critical patent/EP3027866A4/fr
Publication of WO2015017042A1 publication Critical patent/WO2015017042A1/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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16—Control of working fluid flow
    • F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/075—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type controlling flow ratio between flows
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00—Axial-flow pumps
    • F04D19/02—Multi-stage pumps
    • F04D19/028—Layout of fluid flow through the stages
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/545—Ducts
    • F04D29/547—Ducts having a special shape in order to influence fluid flow
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
    • F05D2250/314—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other
    • 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
    • F05D2260/98—Lubrication
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • 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 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 flowpath.
  • 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 in a core flow path axially between the inlet case flow path and the intermediate case flow path.
  • the core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis.
  • a bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.
  • the outer diameter has an outer diameter slope angle relative to the rotational axis along a fluid flow direction of the core flow path of between about 0 degrees and about 15 degrees.
  • a fan is connected to the shaft through a geared architecture.
  • the compressor section is a low pressure compressor.
  • the outer diameter slope angle decreases relative to the rotational axis.
  • the outer diameter slope angle is in the range of about 0 degrees to about 10 degrees.
  • the outer diameter slope angle is in the range of about 5 degrees to about 7 degrees.
  • the outer diameter slope angle is about 6 degrees.
  • a fan is connected to the shaft through a geared architecture.
  • the compressor section is a low pressure compressor.
  • the low pressure compressor comprises at least one variable vane.
  • the low pressure compressor comprises an exit guide vane. The exit guide vane is located in a low pressure compressor outlet section of the core flow path.
  • the low pressure compressor further comprises a low pressure bleed located between a low pressure compressor rotor and the exit guide vane.
  • the low pressure bleed further comprises a bleed trailing edge, and wherein the bleed trailing edge extends into the core flow path beyond the outer diameter of the core flow path.
  • the low pressure compressor is a multi-stage compressor.
  • 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 comprises 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.
  • a 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 includes a second end opposite the first end.
  • the geared architecture includes a sun gear supported on the second end.
  • the shaft includes a hub 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, and the bearing is axially aligned with and radially inward of one of the compressor stages.
  • the compressor section includes a variable vane array.
  • the geared architecture is arranged in the lubrication compartment.
  • 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 contextually illustrates an example core flowpath through a low pressure compressor of the gas turbine engine embodiment of Figure 1.
  • Figure 4 contextually illustrates another example core flowpath through a low pressure compressor of the gas turbine engine embodiment of Figure 1.
  • 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 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 52 and the high pressure turbine 54.
  • 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
  • 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 [(Ta m bie 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.
  • Figure 3 is a sectional view of the gas turbine engine 20 of Figure 1, contextually illustrating a low pressure compressor 44 of the gas turbine engine 20.
  • the core flowpath identified herein as flowpath 220 or core flowpath 220, passes through the low pressure compressor 44 of the gas-turbine engine 20.
  • the low pressure compressor 44 includes multiple rotor 212/vane 214 pairs that serve to drive air through the core flowpath 220.
  • the rotors 212 are connected to an inner shaft 40 via a compressor frame 242. Interspersed between each of the rotors 212 is a vane 214.
  • the vanes 214 are connected to an outer frame 260. Additional stages can be added or removed depending on design constraints via the addition or removal of rotor 212/vane 214 pairs.
  • a variable guide vane 230 is located at an inlet 232 of the low pressure compressor 44.
  • one or more of the vanes 214 could also be a variable vane 230.
  • An exit guide vane 216 is located at a fluid outlet 234 of the low pressure compressor 44. In the illustrated example of Figure 3, the exit guide vane 216 also acts as a vane 214 corresponding to the last rotor 212 of the low pressure compressor 44.
  • the illustrated low pressure compressor 44 is referred to as a three stage compressor as three rotor 212/vane 214 pairs (including vane 216) are included.
  • the core flowpath 220 has an inner diameter 254 and an outer diameter 252 measured with respect to the engine longitudinal axis A. As the core flowpath 220 passes through the low pressure compressor 44, the inner diameter 254 of the core flowpath 220 slopes outward or parallel to relative to the engine central longitudinal axis A away from the engine central longitudinal axis A resulting in an increasing inner diameter 254 as the core flowpath 220 progresses along the direction of fluid flow.
  • the increasing inner diameter 254 may more easily accommodate at least one of the first and second bearings 70, 112, packaged radially inward of the low pressure compressor 44.
  • the outer diameter 252 may slope inward relative to the engine central longitudinal axis A toward the engine central longitudinal axis A to provide a further decreasing cross-sectional area core flowpath 220 that compresses air passing through the low pressure compressor 44.
  • a steeper slope angle of the outer diameter 252, relative to the engine central longitudinal axis A, may result in a greater average tip clearance between the rotor blade 212 and the engine case during flight.
  • the additional tip clearance may increase flow separation in the air flowing through the core flowpath 220.
  • undesirable amounts flow separation can occur when the outer diameter 252 exceeds 15 degrees (absolute value) relative to the engine central longitudinal axis A.
  • Flow separation occurs when the air flow separates from the core flowpath 220 walls.
  • the outer diameter 252 includes a sufficiently low slope angle relative to the engine central longitudinal axis A and then increasing the inner diameter 254, the flow separation resulting from the additional tip clearance may be eliminated (or at least greatly reduced), and the total amount of flow separation may be minimized.
  • a slope angle of the outer diameter 252 is less than about 15 degrees (absolute value), and in some embodiments less than about 10 degrees (absolute value), relative to the engine central longitudinal axis A.
  • the slope angle of the outer diameter 252 is approximately 6 degrees (absolute value) relative to the engine central longitudinal axis A.
  • Figure 4 illustrates an example core flowpath 220.
  • air flow passing through the core flowpath 220 is insufficiently stable.
  • one or more variable guide vanes 230 may be included in the flow path 220.
  • a three stage geared turbofan compressor 44 such as the one illustrated in Figure 3
  • a single variable guide vane 230 can be utilized to sufficiently stabilize the air flow.
  • alternate embodiments, such as those utilizing additional compressor stages may require additional variable guide vanes 230.
  • one or more of the vanes 214 can be the additional variable guide vanes 230.
  • the air flow can be sufficiently stable without the inclusion of a variable guide vane 230, and the variable guide vane 230 can be omitted.
  • the exit guide vane 216 is incorporated into a low pressure compressor outlet 234 section of the core flowpath 220 between the exit of the low pressure compressor 44 and the entrance to the high pressure compressor 52.
  • the low pressure compressor outlet 234 section of the core flowpath 220 is sloped inward (toward the engine central longitudinal axis A). Placing the exit guide vane 216 in the inward sloping low pressure compressor outlet 234 section of the core flowpath 220 cants the exit guide vane 216 and provides space for a low pressure bleed 264.
  • the low pressure bleed 264 allows for dirt, rain and ice to be removed from the compressor 44.
  • the low pressure bleed 264 additionally improves the stability of the fluid flowing through the core flowpath 220.
  • the low pressure bleed 264 is positioned between the last (downstream most) rotor 212 and the exit guide vane 216.
  • a bleed trailing edge 262 of the low pressure bleed 264 can extend inward toward the engine central longitudinal axis A, beyond the outer diameter 252 of the core flowpath 220.
  • the outer diameter of the bleed trailing edge 262 of the low pressure bleed 264 is smaller than the outer diameter 252. Extending the bleed trailing edge 262 inwards allows the bleed 264 to scoop out more of the dirt, rain, ice or other impurities that may enter the core flowpath 220.

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

Abstract

L'invention concerne un moteur de turbine à gaz comprenant un carter central qui renferme un carter d'entrée et un carter intermédiaire fournissant respectivement un trajet d'écoulement de carter d'entrée et un trajet d'écoulement de carter intermédiaire. Un arbre présente un axe rotatif. Un moyeu est supporté fonctionnellement par l'arbre. Un rotor est relié au moyeu et supporte une partie de compresseur. La partie de compresseur est disposée axialement dans un trajet d'écoulement central entre le trajet d'écoulement de carter d'entrée et le trajet d'écoulement de carter intermédiaire. Le trajet d'écoulement central présente un diamètre interne et un diamètre externe. Au moins une partie du diamètre interne présente un angle de pente augmentant par rapport à l'axe rotatif. Un palier est monté sur le moyeu et supporte l'arbre par rapport à l'un des carters d'entrée et intermédiaire.
PCT/US2014/043184 2013-07-31 2014-06-19 Forme de trajet d'écoulement de compresseur basse pression (lpc) avec configuration de palier de moteur de turbine à gaz WO2015017042A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14831790.2A EP3027866A4 (fr) 2013-07-31 2014-06-19 Forme de trajet d'écoulement de compresseur basse pression (lpc) avec configuration de palier de moteur de turbine à gaz

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
US201361860334P 2013-07-31 2013-07-31
US61/860,334 2013-07-31
US14/067,354 2013-10-30
US14/067,354 US9038366B2 (en) 2012-01-31 2013-10-30 LPC flowpath shape with gas turbine engine shaft bearing configuration

Publications (1)

Publication Number Publication Date
WO2015017042A1 true WO2015017042A1 (fr) 2015-02-05

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WO (1) WO2015017042A1 (fr)

Citations (7)

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US5155993A (en) * 1990-04-09 1992-10-20 General Electric Company Apparatus for compressor air extraction
US20050265825A1 (en) * 2004-05-27 2005-12-01 Rolls-Royce Plc Spacing arrangement
US20070087892A1 (en) * 2005-10-19 2007-04-19 General Electric Company Gas turbine engine assembly and methods of assembling same
US20090056306A1 (en) * 2007-08-28 2009-03-05 Suciu Gabriel L Gas turbine engine front architecture
US20120243971A1 (en) * 2006-08-15 2012-09-27 Mccune Michael E Epicyclic gear train
US20120257960A1 (en) * 2009-11-20 2012-10-11 United Technologies Corporation Engine Bearing Support
US20130192198A1 (en) * 2012-01-31 2013-08-01 Lisa I. Brilliant Compressor flowpath

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Publication number Priority date Publication date Assignee Title
US9784181B2 (en) * 2009-11-20 2017-10-10 United Technologies Corporation Gas turbine engine architecture with low pressure compressor hub between high and low rotor thrust bearings
EP3489472B8 (fr) * 2011-10-17 2021-04-07 Raytheon Technologies Corporation Architecture de corps centrale avant de moteur à turbine à gaz
US8402741B1 (en) * 2012-01-31 2013-03-26 United Technologies Corporation Gas turbine engine shaft bearing configuration

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5155993A (en) * 1990-04-09 1992-10-20 General Electric Company Apparatus for compressor air extraction
US20050265825A1 (en) * 2004-05-27 2005-12-01 Rolls-Royce Plc Spacing arrangement
US20070087892A1 (en) * 2005-10-19 2007-04-19 General Electric Company Gas turbine engine assembly and methods of assembling same
US20120243971A1 (en) * 2006-08-15 2012-09-27 Mccune Michael E Epicyclic gear train
US20090056306A1 (en) * 2007-08-28 2009-03-05 Suciu Gabriel L Gas turbine engine front architecture
US20120257960A1 (en) * 2009-11-20 2012-10-11 United Technologies Corporation Engine Bearing Support
US20130192198A1 (en) * 2012-01-31 2013-08-01 Lisa I. Brilliant Compressor flowpath

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
See also references of EP3027866A4 *
WALSH ET AL.: "Gas Turbine Performance", 1998, pages 159 - 177, XP008182806 *

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Publication number Publication date
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