EP2935791A1 - Soufflante carénée légère - Google Patents

Soufflante carénée légère

Info

Publication number
EP2935791A1
EP2935791A1 EP13864578.3A EP13864578A EP2935791A1 EP 2935791 A1 EP2935791 A1 EP 2935791A1 EP 13864578 A EP13864578 A EP 13864578A EP 2935791 A1 EP2935791 A1 EP 2935791A1
Authority
EP
European Patent Office
Prior art keywords
fan
gas turbine
section
turbine engine
shroud
Prior art date
Legal status (The legal status 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 status listed.)
Withdrawn
Application number
EP13864578.3A
Other languages
German (de)
English (en)
Other versions
EP2935791A4 (fr
Inventor
Frederick M. Schwarz
Michael A. Weisse
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
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
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2935791A1 publication Critical patent/EP2935791A1/fr
Publication of EP2935791A4 publication Critical patent/EP2935791A4/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • 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
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/325—Rotors specially for elastic fluids for axial flow pumps for axial flow fans
    • F04D29/326—Rotors specially for elastic fluids for axial flow pumps for axial flow fans comprising a rotating shroud
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/004—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids by varying driving speed
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0261—Surge control by varying driving speed
    • 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
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • 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
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • 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
    • Y02B—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO BUILDINGS, e.g. HOUSING, HOUSE APPLIANCES OR RELATED END-USER APPLICATIONS
    • Y02B30/00—Energy efficient heating, ventilation or air conditioning [HVAC]
    • Y02B30/70—Efficient control or regulation technologies, e.g. for control of refrigerant flow, motor or heating
    • 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
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49229—Prime mover or fluid pump making

Definitions

  • a gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
  • the fan section includes multiple airfoils disposed circumferentially about an engine longitudinal centerline axis. At certain aircraft operating conditions, these airfoils may experience self-induced oscillations, such as flutter. These self-induced oscillations may become severe enough to fracture the airfoil.
  • One means of preventing such a fracture is to increase the chord width of the fan blades. However, this approach increases the overall weight of the engine and the rotating mass. Accordingly, it is desirable to develop an improved gas turbine engine design that will reduce flutter of the airfoils and decrease the weight of the engine.
  • a gas turbine engine includes, among other things, a fan section including a fan with a plurality of fan blades rotatable about an axis and a shroud and a speed change device in communication with the fan.
  • the speed change device includes a geared architecture driven by a turbine section for rotating the fan about the axis.
  • the shroud is located at radially outer ends of the plurality of fan blades.
  • the gas turbine engine includes a low pressure turbine with at least three stages and no more than six stages.
  • the gas turbine engine includes a fixed area nozzle in communication with the fan section.
  • the gas turbine engine includes a variable area nozzle in communication with the fan section.
  • the gas turbine engine includes a seal between the shroud and a fan case.
  • the seal includes at least one knife edge.
  • the seal includes a honeycomb structure.
  • the seal includes a rubber member on a radially inner surface of the fan case.
  • the shroud is made of a fiber composite.
  • a method of assembling a gas turbine engine includes, among other things, positioning a turbine section in communication with a shaft.
  • a speed change device is positioned in communication with the shaft.
  • a fan section is positioned in communication with the speed change device.
  • the fan section includes a plurality of fan blades and a shroud.
  • the speed change device includes a geared architecture.
  • the shroud is located at the radially outer ends of the plurality of fan blades.
  • the turbine section is a low pressure turbine.
  • the turbine section includes at least three stages and no more than six stages.
  • a fixed area fan nozzle is positioned in communication with the fan section.
  • a method of operating a gas turbine engine includes, among other things, rotating a fan section including a shroud at a first speed and rotating a turbine section at a second speed. The first speed is different from the second speed.
  • a speed change device is in mechanical communication with the fan section and the turbine section.
  • the turbine section is a low pressure turbine.
  • Figure 1 is a schematic view of an example gas turbine engine.
  • Figure 2 is a perspective view of a shrouded fan.
  • Figure 3 is a partial cross- sectional view of a fixed area fan nozzle.
  • Figure 4 is a partial cross- sectional view of a variable area fan nozzle.
  • FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmenter section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26.
  • air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
  • turbofan gas turbine engine depicts a turbofan gas turbine engine
  • the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three- spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
  • the example 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.
  • the low speed spool 30 generally includes an inner shaft 40 that connects a fan 62 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46.
  • the inner shaft 40 drives the fan 62 through a speed change device, such as a geared architecture 48, to drive the fan 62 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 (or second) compressor section 52 and a high pressure (or second) turbine section 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54.
  • the high pressure turbine 54 includes only a single stage.
  • a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure” compressor or turbine.
  • the example low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the low pressure turbine 46 includes at least three stages and no more than 6 stages. In another non- limiting embodiment, the low pressure turbine 46 includes at least three stages and no more than 4 stages.
  • a mid-turbine frame 58 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 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.
  • Air flowing through the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid- turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
  • the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
  • the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10).
  • the example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
  • the gas turbine engine 20 includes a bypass ratio greater than about ten (10: 1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. [0039] 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
  • 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 1.50. In another non- limiting embodiment the low fan pressure ratio 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 [(Tram °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.
  • the example gas turbine engine includes the fan 62 that comprises in one non-limiting embodiment less than about 26 fan blades 42. In another non-limiting embodiment, the fan section 22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about 3 turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
  • the example gas turbine engine 20 includes fan blades 42 that extend from a central disk 64 on a radially inner end to a shroud 66 on a radially outer end.
  • the shroud 66 is comprised of a fiber composite, such as a woven fiber composite or a wound fiber composite. Due to the geared architecture 48, the fan 62 rotates at a slower speed than the low pressure turbine 46. Because the fan 62 has a lower rotational speed, the fan blade tip velocity decreases and the aerodynamic losses that would normally accompany a shrouded fan are reduced.
  • the gas turbine engine 20 generates a similar amount of thrust as a gas turbine engine with a fan section that rotates at the same speed as the low pressure turbine by increasing the length and number of fan blades 42.
  • the fan 62 accommodates more fan blades 42 by decreasing the chord width of the fan blades 42 to allow for more fan blades 42.
  • Increasing the length and number of fan blades 42 and decreasing the chord width of the individual fan blades 42 will decrease the overall weight of the gas turbine engine 20 as well as the rotating mass.
  • the fan section 22 includes an example fan case 74.
  • An outer surface 68 of the shroud 66 includes at least one knife edge 70 that form a seal between the outer surface 68 of the shroud 66 and an inner surface 72 of the fan case 74 to prevent air leakage between the shroud 66 and the fan case 74.
  • the at least one knife edge 70 includes three knife edges that extend around the outer surface 68 of the shroud 66 ( Figure 2).
  • the inner surface 72 of the fan case 74 includes an outer air seal 76, such as a rubber member, to prevent air leakage between the shroud 66 and the fan case 74.
  • the outer air seal 76 accommodates for expansion of the fan 62 without causing damage to the fan case 74 because of the elastic nature of the outer air seal 76.
  • the shroud 66 includes a honeycomb structure 78 on the outer surface 68 for forming an additional seal between the shroud 66 and the fan case 74 by creating a disturbance in the air flowing between the shroud 66 and the fan case 74.
  • the fan nozzle 65 includes a fixed area fan nozzle such that the exit area for the fan section 22 is fixed during operation of the gas turbine engine 20 ( Figure 3). Eliminating a variable area fan nozzle from the gas turbine engine 20 provides a significant weight loss over convention gas turbine engines with variable area fan nozzles. A variable area fan nozzle can be eliminated from the gas turbine engine 20 because of the gas turbine engine's 20 ability to prevent flutter through the use of the shroud 66 and the lower rotational speed of the fan 62 due to the geared architecture 48.
  • the fan nozzle 65' includes a variable area fan nozzle such that the exit area of the fan section 22 is varied during operation of the gas turbine engine 20 ( Figure 4). Increasing the exit area of the fan section will prevent flutter of the fan blades 42 from occurring by decreasing the pressure downstream of the fan blades 42.

Landscapes

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

Abstract

Un moteur à turbine à gaz selon un mode de réalisation de la présente invention donné à titre d'exemple comprend, entre autres, une section de soufflante qui comprend une soufflante pourvue d'une pluralité d'aubes pouvant tourner autour d'un axe et un carénage et un dispositif de changement de vitesse en communication avec la soufflante.
EP13864578.3A 2012-12-19 2013-12-16 Soufflante carénée légère Withdrawn EP2935791A4 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/719,859 US20140212261A1 (en) 2012-12-19 2012-12-19 Lightweight shrouded fan
PCT/US2013/075260 WO2014099713A1 (fr) 2012-12-19 2013-12-16 Soufflante carénée légère

Publications (2)

Publication Number Publication Date
EP2935791A1 true EP2935791A1 (fr) 2015-10-28
EP2935791A4 EP2935791A4 (fr) 2016-01-20

Family

ID=50979065

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13864578.3A Withdrawn EP2935791A4 (fr) 2012-12-19 2013-12-16 Soufflante carénée légère

Country Status (3)

Country Link
US (1) US20140212261A1 (fr)
EP (1) EP2935791A4 (fr)
WO (1) WO2014099713A1 (fr)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150285259A1 (en) * 2014-04-05 2015-10-08 Arthur John Wennerstrom Filament-Wound Tip-Shrouded Axial Compressor or Fan Rotor System
US10344711B2 (en) * 2016-01-11 2019-07-09 Rolls-Royce Corporation System and method of alleviating blade flutter
US10465539B2 (en) * 2017-08-04 2019-11-05 Pratt & Whitney Canada Corp. Rotor casing
US10724535B2 (en) 2017-11-14 2020-07-28 Raytheon Technologies Corporation Fan assembly of a gas turbine engine with a tip shroud

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Publication number Priority date Publication date Assignee Title
US3556675A (en) * 1969-01-29 1971-01-19 Gen Electric Turbomachinery rotor with integral shroud
US3580692A (en) * 1969-07-18 1971-05-25 United Aircraft Corp Seal construction
GB1432994A (en) * 1973-05-02 1976-04-22 Rolls Royce Compressor for gas turbine engines
US4239452A (en) * 1978-06-26 1980-12-16 United Technologies Corporation Blade tip shroud for a compression stage of a gas turbine engine
FR2761412B1 (fr) * 1997-03-27 1999-04-30 Snecma Groupe turbopropulseur double corps a regulation isodrome
US6203273B1 (en) * 1998-12-22 2001-03-20 United Technologies Corporation Rotary machine
US6881036B2 (en) * 2002-09-03 2005-04-19 United Technologies Corporation Composite integrally bladed rotor
US6814541B2 (en) * 2002-10-07 2004-11-09 General Electric Company Jet aircraft fan case containment design
DE602004020124D1 (de) * 2004-12-08 2009-04-30 Volvo Aero Corp Rad für eine rotationsströmungsmaschine
US7513103B2 (en) * 2005-10-19 2009-04-07 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
US7254997B1 (en) * 2006-10-12 2007-08-14 David Hui Anti-steal tire pressure monitoring apparatus
US8844265B2 (en) * 2007-08-01 2014-09-30 United Technologies Corporation Turbine section of high bypass turbofan
GB2483495B (en) * 2010-09-10 2013-02-13 Magnaparva Space Ltd Mounting of rotor blades
US20140169972A1 (en) * 2012-12-17 2014-06-19 United Technologies Corporation Fan with integral shroud

Also Published As

Publication number Publication date
WO2014099713A1 (fr) 2014-06-26
EP2935791A4 (fr) 2016-01-20
US20140212261A1 (en) 2014-07-31

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