US20130108432A1 - Airflow-straightening structure for the nacelle of an aircraft engine - Google Patents

Airflow-straightening structure for the nacelle of an aircraft engine Download PDF

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Publication number
US20130108432A1
US20130108432A1 US13/718,562 US201213718562A US2013108432A1 US 20130108432 A1 US20130108432 A1 US 20130108432A1 US 201213718562 A US201213718562 A US 201213718562A US 2013108432 A1 US2013108432 A1 US 2013108432A1
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US
United States
Prior art keywords
hoop
airflow
clevis
straightening
hub
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.)
Abandoned
Application number
US13/718,562
Inventor
Bertrand Desjoyeaux
Nicolas Dezeustre
Wouter Balk
François Gallet
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.)
Safran Aircraft Engines SAS
Safran Nacelles SAS
Original Assignee
Aircelle SA
SNECMA SAS
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 Aircelle SA, SNECMA SAS filed Critical Aircelle SA
Publication of US20130108432A1 publication Critical patent/US20130108432A1/en
Assigned to SNECMA, AIRCELLE reassignment SNECMA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BALK, WOUTER, GALLET, FRANCOIS, DESJOYEAUX, BERTRAND, DEZEUSTRE, NICOLAS
Abandoned legal-status Critical Current

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Classifications

    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
    • 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
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • F01D5/34—Rotor-blade aggregates of unitary construction, e.g. formed of sheet laminae
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • 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
    • F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep
    • 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
    • F02K3/06—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 with front fan
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/14—Casings or housings protecting or supporting assemblies within
    • 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
    • F05D2250/50—Inlet or outlet
    • F05D2250/51—Inlet
    • 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
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure relates to an airflow-straightening structure for an aircraft engine.
  • a nacelle having axis A for a dual-flow (turbojet) engine traditionally comprises an outer structure 1 having an upstream portion 3 forming an air intake, an intermediate portion 5 whereof the inner skin 6 forms a case for the fan 7 of the engine, and a downstream portion 9 that may incorporate thrust reversal means.
  • This nacelle also includes an inner structure 11 having a fairing 13 for the engine 15 .
  • the outer structure 1 defines, with the inner structure 11 , an annular air duct 17 , often called “cold air duct,” as opposed to the hot air created by the engine 15 .
  • the fan 7 essentially consists of a propeller provided with blades 19 , which are rotatably mounted on a stationary hub 21 connected to the fan case 6 by a plurality of stationary arms 25 , which may for example be distributed at 120 degree intervals.
  • airflow-straightening vanes 23 also called OGV (“Outlet Guide Vanes”), which make it possible to straighten the cold air flow created by the fan 7 .
  • the present disclosure provides an airflow-straightening structure for an aircraft engine, comprising:
  • the structure being remarkable in that at least two of the elements chosen from the group consisting of said hoop, said plurality of vanes, said hub and said clevis are formed in one piece, that is to say without any assembly operation,
  • Having at least two of the aforementioned elements (hoop, vanes, hub, clevis) made in a single piece makes it possible to limit the number of assembly operations to be carried out during assembly of the nacelle, and having at least one of those four elements made from a composite material enables savings in terms of weight.
  • the present disclosure also relates to a nacelle for an aircraft engine, remarkable in that it comprises an airflow-straightening structure as described above.
  • FIG. 1 is a longitudinal half-section view of a nacelle and engine assembly of the prior art, described in the preamble of this description;
  • FIG. 2 is a perspective view of an airflow-straightening structure according to the present disclosure.
  • FIG. 3 is a detailed cross-sectional view of an engine equipped with an airflow-straightening structure according to one particular form of the present disclosure, and the associated nacelle portion.
  • FIG. 2 shows that the airflow-straightening structure according to the invention may comprise a plurality of airflow-straightening vanes 23 extending between a radially inner wheel 27 and a radially outer wheel 29 , these vanes and wheels thus forming a flow-straightening grid.
  • the wheel with the larger diameter 29 is designed to be fixed or integrated inside a hoop 31 that is part of the inner skin of the intermediate portion 5 of the outer structure 1 of the engine.
  • the wheel with the smaller diameter 27 bears the stationary hub 21 on which the fan 7 is rotatably mounted.
  • a clevis 33 for suspending the nacelle from a support strut secured to the structure of an aircraft: this clevis makes it possible to connect the nacelle and its associated engine 15 to the aircraft.
  • the airflow-straightening vanes 23 and the wheels 27 , 29 are formed in a single piece.
  • the hub 21 on the one hand and the assembly formed by the hoop 31 and the clevis 33 on the other hand are also formed in a single piece.
  • At least some of said members are made from a composite material, for example by weaving.
  • the aforementioned set of members i.e. the hub 21 , the vanes 23 , the wheels 27 , 29 , the hoop 31 and the yoke 33 , are formed in a single piece, and from a composite material.
  • FIG. 3 shows a detailed view of the upstream portion of a nacelle and part of the engine which are equipped with an airflow-straightening structure according to the invention, and shows that the hoop 31 can incorporate a fan case 6 and the air intake shroud 35 .
  • the air intake shroud 35 is a substantially cylindrical piece forming the inner skin of the air intake 3 , on which sound absorption panels 37 are attached that can typically have a honeycomb structure, so as to form Helmholtz resonators.
  • the hoop 31 can also incorporate, in the downstream portion thereof, an extension 37 forming a cascade edge, i.e. defining one of the walls of the channel making it possible to guide the cold airflow toward the outside of the nacelle during implementation of the thrust reversal means (not shown).
  • the additional integration of the fan case 6 , the shroud 35 and the cascade edge 37 in the flow-straightening structure according to the invention also allows increased structural simplicity, as well as weight savings when the set of parts is made in a single piece and from a composite material.

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

Abstract

An airflow-straightening structure for an aircraft engine is provided that includes a hoop inside which there are arranged a plurality of flow-straightening vanes bearing a hub of a fan. Also includes is a clevis for connecting to a suspension strut, the clevis being fixed to said hoop. The structure has at least two elements chosen from the group of the hoop, the plurality of vanes, the hub and the clevis, which are formed in one piece, that is to say without any assembly operation. Additionally, at least one of the elements chosen from the group is at least partially formed of a composite material.

Description

    CROSS-REFERENCE TO RELATED APPLICATIONS
  • This application is a continuation of International Application No. PCT/FR2011/051379 filed on Jun. 16, 2011, which claims the benefit of FR 10/54852, filed on Jun. 18, 2010. The disclosures of the above applications are incorporated herein by reference.
    • FIELD
  • The present disclosure relates to an airflow-straightening structure for an aircraft engine.
    • BACKGROUND
  • The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
  • As is known in itself, and shown in the appended FIG. 1, a nacelle having axis A for a dual-flow (turbojet) engine traditionally comprises an outer structure 1 having an upstream portion 3 forming an air intake, an intermediate portion 5 whereof the inner skin 6 forms a case for the fan 7 of the engine, and a downstream portion 9 that may incorporate thrust reversal means.
  • This nacelle also includes an inner structure 11 having a fairing 13 for the engine 15.
  • The outer structure 1 defines, with the inner structure 11, an annular air duct 17, often called “cold air duct,” as opposed to the hot air created by the engine 15.
  • The fan 7 essentially consists of a propeller provided with blades 19, which are rotatably mounted on a stationary hub 21 connected to the fan case 6 by a plurality of stationary arms 25, which may for example be distributed at 120 degree intervals.
  • Upstream of these stationary arms are airflow-straightening vanes 23, also called OGV (“Outlet Guide Vanes”), which make it possible to straighten the cold air flow created by the fan 7.
    • SUMMARY
  • The present disclosure provides an airflow-straightening structure for an aircraft engine, comprising:
      • a hoop inside which there are arranged a plurality of flow-straightening vanes bearing a hub of a fan, and
      • a clevis for connecting to a suspension strut, said clevis being fixed to said hoop,
  • the structure being remarkable in that at least two of the elements chosen from the group consisting of said hoop, said plurality of vanes, said hub and said clevis are formed in one piece, that is to say without any assembly operation,
  • and for the fact that at least one of the elements chosen from the group consisting of said hoop, said plurality of vanes, said hub and said clevis is at least partially formed of composite material.
  • Having at least two of the aforementioned elements (hoop, vanes, hub, clevis) made in a single piece makes it possible to limit the number of assembly operations to be carried out during assembly of the nacelle, and having at least one of those four elements made from a composite material enables savings in terms of weight.
  • According to other optional features of this structure according to the invention:
      • all of said elements are made in a single piece and from a composite material: in this way, the ease of assembly and weight reduction are optimized;
      • the hoop incorporates the case of said fan;
      • the hoop incorporates an air intake shroud for the nacelle;
      • the hoop incorporates a thrust reverser cascade edge: these various arrangements make it possible to still further reduce the number of parts.
  • The present disclosure also relates to a nacelle for an aircraft engine, remarkable in that it comprises an airflow-straightening structure as described above.
  • Further areas of applicability will become apparent from the description provided herein. It should be understood that the description and specific examples are intended for purposes of illustration only and are not intended to limit the scope of the present disclosure.
    • DRAWINGS
  • In order that the disclosure may be well understood, there will now be described various forms thereof, given by way of example, reference being made to the accompanying drawings, in which:
  • FIG. 1 is a longitudinal half-section view of a nacelle and engine assembly of the prior art, described in the preamble of this description;
  • FIG. 2 is a perspective view of an airflow-straightening structure according to the present disclosure; and
  • FIG. 3 is a detailed cross-sectional view of an engine equipped with an airflow-straightening structure according to one particular form of the present disclosure, and the associated nacelle portion.
    •
  • In all of the figures, identical or similar references designate identical or similar members or sets of members.
  • It will also be noted that a three-axis reference has been provided in these figures showing the X, Y and Z axes. These three axes respectively represent the longitudinal, transverse and vertical directions of the engine when it is installed on an aircraft.
  • The drawings described herein are for illustration purposes only and are not intended to limit the scope of the present disclosure in any way.
    • DETAILED DESCRIPTION
  • The following description is merely exemplary in nature and is not intended to limit the present disclosure, application, or uses. It should be understood that throughout the drawings, corresponding reference numerals indicate like or corresponding parts and features.
  • FIG. 2 shows that the airflow-straightening structure according to the invention may comprise a plurality of airflow-straightening vanes 23 extending between a radially inner wheel 27 and a radially outer wheel 29, these vanes and wheels thus forming a flow-straightening grid.
  • The wheel with the larger diameter 29 is designed to be fixed or integrated inside a hoop 31 that is part of the inner skin of the intermediate portion 5 of the outer structure 1 of the engine.
  • The wheel with the smaller diameter 27 bears the stationary hub 21 on which the fan 7 is rotatably mounted.
  • Fixed on the upper portion of the hoop 31 is a clevis 33 for suspending the nacelle from a support strut secured to the structure of an aircraft: this clevis makes it possible to connect the nacelle and its associated engine 15 to the aircraft.
  • In one form, the airflow-straightening vanes 23 and the wheels 27, 29 are formed in a single piece.
  • Also, the hub 21 on the one hand and the assembly formed by the hoop 31 and the clevis 33 on the other hand are also formed in a single piece.
  • At least some of said members are made from a composite material, for example by weaving.
  • Still more preferably, the aforementioned set of members, i.e. the hub 21, the vanes 23, the wheels 27, 29, the hoop 31 and the yoke 33, are formed in a single piece, and from a composite material.
  • This makes it possible to obtain a multi-functional piece with an overall weight much lower than the set of parts it replaces, and not requiring any assembly operation.
  • As can be understood from the preceding, the fact that the hub 21 of the fan 7 is borne directly by the airflow-straightening vanes 23 makes it possible to do away with the support arms for that hub, which are present in the nacelles according to the prior art: savings are thus procured in terms of structural simplicity and weight.
  • FIG. 3 shows a detailed view of the upstream portion of a nacelle and part of the engine which are equipped with an airflow-straightening structure according to the invention, and shows that the hoop 31 can incorporate a fan case 6 and the air intake shroud 35.
  • As a reminder, the air intake shroud 35 is a substantially cylindrical piece forming the inner skin of the air intake 3, on which sound absorption panels 37 are attached that can typically have a honeycomb structure, so as to form Helmholtz resonators.
  • The hoop 31 can also incorporate, in the downstream portion thereof, an extension 37 forming a cascade edge, i.e. defining one of the walls of the channel making it possible to guide the cold airflow toward the outside of the nacelle during implementation of the thrust reversal means (not shown).
  • As will be understood, the additional integration of the fan case 6, the shroud 35 and the cascade edge 37 in the flow-straightening structure according to the invention also allows increased structural simplicity, as well as weight savings when the set of parts is made in a single piece and from a composite material.
  • Of course, the present disclosure is in no way limited to the embodiments described and shown, which have been provided solely as examples.

Claims (6)

What is claimed is:
1 An airflow-straightening structure for an aircraft engine, comprising:
a hoop inside which there are arranged a plurality of flow-straightening vanes bearing a hub of a fan; and
a clevis for connecting to a suspension strut, said clevis being fixed to said hoop, the structure being characterized in that at least two elements chosen from a group consisting of said hoop, said plurality of vanes, said hub and said clevis are formed in one piece, that is to say without any assembly operation,
and in that at least one of the elements chosen from the group consisting of said hoop, said plurality of vanes, said hub and said clevis is at least partially formed of a composite material.
2. The structure according to claim 1, characterized in that all of said elements are made in a single piece and from a composite material.
3. The structure according to claim 1, characterized in that said hoop incorporates the case of said fan.
4. The structure according to claim 1, characterized in that said hoop incorporates an air intake shroud.
5. The structure according to claim 1, characterized in that said hoop incorporates a thrust reverser cascade edge.
6. A nacelle for an aircraft engine, characterized in that it comprises an airflow-straightening structure according to claim 1.
US13/718,562 2010-06-18 2012-12-18 Airflow-straightening structure for the nacelle of an aircraft engine Abandoned US20130108432A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1054852A FR2961555B1 (en) 2010-06-18 2010-06-18 AIR FLOW RECTIFYING STRUCTURE FOR AN AIRCRAFT ENGINE NACELLE
FR10/54852 2010-06-18
PCT/FR2011/051379 WO2011157962A1 (en) 2010-06-18 2011-06-16 Airflow-straightening structure for the nacelle of an aircraft engine

Related Parent Applications (1)

Application Number Title Priority Date Filing Date
PCT/FR2011/051379 Continuation WO2011157962A1 (en) 2010-06-18 2011-06-16 Airflow-straightening structure for the nacelle of an aircraft engine

Publications (1)

Publication Number Publication Date
US20130108432A1 true US20130108432A1 (en) 2013-05-02

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US (1) US20130108432A1 (en)
EP (1) EP2582580A1 (en)
CN (1) CN102971217A (en)
BR (1) BR112012030323A2 (en)
CA (1) CA2802590A1 (en)
FR (1) FR2961555B1 (en)
RU (1) RU2013101809A (en)
WO (1) WO2011157962A1 (en)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150132110A1 (en) * 2012-06-15 2015-05-14 United Technologies Corporation High durability turbine exhaust case
USD732656S1 (en) * 2013-07-25 2015-06-23 Asustek Computer Inc. Fan blade
US20160108757A1 (en) * 2014-10-16 2016-04-21 Honeywell International Inc. Integrated outer flowpath ducting and front frame system for use in a turbofan engine and method for making same
EP3734017A1 (en) * 2019-05-03 2020-11-04 Raytheon Technologies Corporation Gas turbine engine with fan case having integrated stator vanes

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2900941B1 (en) * 2012-09-26 2016-12-14 United Technologies Corporation Combined high pressure turbine case and turbine intermediate case

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2857093A (en) * 1954-12-02 1958-10-21 Cincinnati Testing & Res Lab Stator casing and blade assembly
US6145300A (en) * 1998-07-09 2000-11-14 Pratt & Whitney Canada Corp. Integrated fan / low pressure compressor rotor for gas turbine engine
US6312215B1 (en) * 2000-02-15 2001-11-06 United Technologies Corporation Turbine engine windmilling brake
US6843449B1 (en) * 2004-02-09 2005-01-18 General Electric Company Fail-safe aircraft engine mounting system
US20060277895A1 (en) * 2005-05-11 2006-12-14 Thornock Russel L Aircraft systems including cascade thrust reversers
US20090120058A1 (en) * 2004-12-01 2009-05-14 United Technologies Corporation Tip Turbine Engine Integral Fan, Combustor, and Turbine Case

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2161109B (en) * 1984-07-07 1988-12-21 Rolls Royce Integral bladed member
GB2161110B (en) * 1984-07-07 1988-03-23 Rolls Royce An annular bladed member having an integral shroud and a method of manufacture thereof
JP3831265B2 (en) * 2002-01-21 2006-10-11 本田技研工業株式会社 Method for manufacturing stationary blade structure
US6821087B2 (en) * 2002-01-21 2004-11-23 Honda Giken Kogyo Kabushiki Kaisha Flow-rectifying member and its unit and method for producing flow-rectifying member
US7370467B2 (en) * 2003-07-29 2008-05-13 Pratt & Whitney Canada Corp. Turbofan case and method of making
WO2009157817A1 (en) * 2008-06-26 2009-12-30 Volvo Aero Corporation Vane assembly and method of fabricating, and a turbo-machine with such vane assembly

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2857093A (en) * 1954-12-02 1958-10-21 Cincinnati Testing & Res Lab Stator casing and blade assembly
US6145300A (en) * 1998-07-09 2000-11-14 Pratt & Whitney Canada Corp. Integrated fan / low pressure compressor rotor for gas turbine engine
US6312215B1 (en) * 2000-02-15 2001-11-06 United Technologies Corporation Turbine engine windmilling brake
US6843449B1 (en) * 2004-02-09 2005-01-18 General Electric Company Fail-safe aircraft engine mounting system
US20090120058A1 (en) * 2004-12-01 2009-05-14 United Technologies Corporation Tip Turbine Engine Integral Fan, Combustor, and Turbine Case
US20060277895A1 (en) * 2005-05-11 2006-12-14 Thornock Russel L Aircraft systems including cascade thrust reversers

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150132110A1 (en) * 2012-06-15 2015-05-14 United Technologies Corporation High durability turbine exhaust case
US10036276B2 (en) * 2012-06-15 2018-07-31 United Technologies Corporation High durability turbine exhaust case
USD732656S1 (en) * 2013-07-25 2015-06-23 Asustek Computer Inc. Fan blade
USD744085S1 (en) 2013-07-25 2015-11-24 Asustek Computer Inc. Fan blade
US20160108757A1 (en) * 2014-10-16 2016-04-21 Honeywell International Inc. Integrated outer flowpath ducting and front frame system for use in a turbofan engine and method for making same
US9816396B2 (en) * 2014-10-16 2017-11-14 Honeywell International Inc. Integrated outer flowpath ducting and front frame system for use in a turbofan engine and method for making same
EP3734017A1 (en) * 2019-05-03 2020-11-04 Raytheon Technologies Corporation Gas turbine engine with fan case having integrated stator vanes

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RU2013101809A (en) 2014-07-27
WO2011157962A1 (en) 2011-12-22
CA2802590A1 (en) 2011-12-22
FR2961555A1 (en) 2011-12-23
EP2582580A1 (en) 2013-04-24
FR2961555B1 (en) 2014-04-18
BR112012030323A2 (en) 2016-08-09

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