EP1059420A1 - Gehäuse für einen Hochdruckkompressor - Google Patents

Gehäuse für einen Hochdruckkompressor Download PDF

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Publication number
EP1059420A1
EP1059420A1 EP00401609A EP00401609A EP1059420A1 EP 1059420 A1 EP1059420 A1 EP 1059420A1 EP 00401609 A EP00401609 A EP 00401609A EP 00401609 A EP00401609 A EP 00401609A EP 1059420 A1 EP1059420 A1 EP 1059420A1
Authority
EP
European Patent Office
Prior art keywords
ventilation
ferrule
sectors
rings
stator according
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.)
Granted
Application number
EP00401609A
Other languages
English (en)
French (fr)
Other versions
EP1059420B1 (de
Inventor
Pascal Gérard Gervais
Pascal Michel Daniel Lejeune
Carmen Miraucourt
Jacky Serge Naudet
Patrice Suet
Monique Andrée Thore
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
Original Assignee
SNECMA Moteurs SA
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 SNECMA Moteurs SA filed Critical SNECMA Moteurs SA
Publication of EP1059420A1 publication Critical patent/EP1059420A1/de
Application granted granted Critical
Publication of EP1059420B1 publication Critical patent/EP1059420B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
    • 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

Definitions

  • the subject of this invention is a stator heterogeneous structure likely to apply in particular to high pressure compressors of gas turbines.
  • the rotor and stator structure of gas turbine is often cooled or ventilated by the air taken from the flow which flows through the machine.
  • air taken from the flow which flows through the machine There are even double breakdowns associated with double samples, where a breakdown of a downstream part of the stator and rotor follows a first ventilation of the stator and rotor carried out further upstream.
  • the air taken for ventilation in downstream comes from a part of the machine where it has already been compressed, which made it much warmer than the air from the upstream ventilation.
  • the invention consists of a compressor stator fitted with a upstream ventilation and downstream ventilation warmer than upstream ventilation and including a ferrule defining a gas flow stream, characterized in that it comprises a first portion ferrule, subject to upstream ventilation, to continuous ring structure on a circumference and in a first material, and a second portion of ferrule, subjected to downstream ventilation, with structure formed of juxtaposed angular sectors and in one second material having a coefficient of expansion larger than the first material.
  • the first and second materials can be chosen, respectively, from materials with a lower coefficient of expansion such as TA6V and titanium alloys, INC0909, intermetallic of the TiAl type, having an average coefficient of linear expansion of less than 10.10 -6 m per degree; and among materials with a greater coefficient of expansion such as nickel-based alloys of the type INC0718, RENE77 and derivatives, having an average coefficient of linear expansion close to 15.10 -6 m per degree.
  • materials with a lower coefficient of expansion such as TA6V and titanium alloys, INC0909, intermetallic of the TiAl type, having an average coefficient of linear expansion of less than 10.10 -6 m per degree
  • materials with a greater coefficient of expansion such as nickel-based alloys of the type INC0718, RENE77 and derivatives, having an average coefficient of linear expansion close to 15.10 -6 m per degree.
  • figure 1 is an overview of a high pressure compressor a gas turbine
  • Figure 2 is an enlarged view the downstream part of the stator of this compressor
  • Figure 2A a similar view of another embodiment possible of the invention
  • Figures 3 and 4 are two sections of the upstream part and the downstream part of the compressor
  • Figure 5 is an enlarged view of the upstream part of the compressor.
  • a high pressure compressor such as that of FIG. 1 comprises a central rotor 1 driven by a line of trees 2 and composed of a envelope 3 of tapered shape composed of rings 4 juxtaposed and separated by 5 discs at right stages of movable blades 6.
  • a stator 7 surrounds the rotor 1 and comprises, in internal lining of a carcass 8, a portion 9 to which the invention relates and which consists of a support casing 10 and a ferrule 11 supported by the casing 10, facing the rotor 1 and which is used to define an annular vein 12 of gas flow in which the movable blade stages 6 and blade stages stationary 13 flow straightening, which are attached to the ferrule 11 and alternate with the previously mentioned floors.
  • tips of stationary vanes 13, located in front the casing 3 of the rotor 1, carry rings of connection 14 furnished with circular bands of said material abradable 15, formed of a honeycomb structure or more generally of easy erosion, which is dug by facing ribs 16 erected on the casing 3 and which form with it a seal at labyrinth.
  • the tips of the moving blades 6 are free of all equipment and end up close of the shell 11.
  • the internal portion 9 of the stator 7 has discontinuities, which are openings of air sampling from vein 12, noted by references 17, 18 and which give in rooms 19 and 20 respectively established between portion 9 and carcass 8 and through which the air taken from the vein 12 to ventilate in particular the casing 10 and subject it to temperature and expansion determined thermal.
  • the inside of rotor 1 is also ventilated, firstly through a bore 21 of the casing 3 located upstream of the rotor 1 and by which fresh air at about the same temperature as whoever enters room 19 is sucked out, then by another bore 22 of the casing 3, substantially at the right of the second opening 18.
  • the rooms 19 and 20 divide stator 7 into two zones of ventilation, in front of which they extend respectively and which are located on both sides of the opening 19 for entry into the downstream chamber 20, which divide portion 9 in half.
  • Sectors 23 and 23 'adjacent are joined by flexible tabs 24 sealing, extending in grooves longitudinal edges of the sectors and joining by their ends 25, between circles of sectors 23 and 23 'consecutive; and other tabs 26 flexible established in grooves purely or obliquely radial from the edges of sectors 23 and 23 ', and extending from the first tabs 24 to the casing 10.
  • This arrangement effectively prevents gases, very hot in this place, from the vein 12 to leak between the sectors 23 and 23 'to reach the housing 10 and risk damaging it.
  • the tabs 24 and 26 isolate empty volumes 27 (which can be filled with insulation at the heat) that appear between each of the circles of sectors 23 and 23 'and associated rings 28 of the casing 10.
  • each of them includes a rear lip 31, projecting towards inside and back, and which is enclosed between a lip 32 of one of the rings 28, located radially outwards, and a lip 33 or 33 ' pointing forward and drawn either in front of sectors 23, either at the front of the ring 28 located on further downstream; and sectors 23 and 23 'include another outer lip 34 at the front, which cooperates with the lips 33 to grip the lips together 31 and 32 directed towards the rear.
  • Sectors 23 ' differ in that they only understand one lip single at the front, bearing the reference 35 and oriented backwards, and which is housed in a groove 36 of the ring 28 located furthest forward.
  • This mode is simpler than a fashion inspired by more traditional ring attachment designs ferrule, illustrated in FIG. 2A, where the lips 31 and 32 are joined by separate seals 37 with a cross-section clip and where the ferrule elements include a relatively high rib 38 ending in a lip 39 facing forward and housed in a groove of the adjacent ring; however, it is possible to adopt this less favorable conception if desired.
  • Of 50 stud systems allow in all the cases of linking sectors 23 and 23 'to the rings 28 in angular direction; many achievements are within the reach of the skilled person.
  • the mode of connection of sectors 23 and 23 'to rings 28 is quite flexible and absorbs deformations without receiving strong constraints.
  • the rings 28 are preferably continuous on the circumference for give a simpler structure and better mechanical resistance.
  • rotor 1 in the same material opposite the rings 28 of the stator 7.
  • a nickel-based alloy, type INCO718, with high coefficient of expansion can be used for this downstream part of the compressor.
  • the housing 10 is at this place composed of 40 rings, united between them by bolts 42 enclosing flanges 41 which complete, as well as the carcass 8, in the manner of rings 28; but these 40 rings still include protrusions 43 and 43 'radially inside, which lead to the air flow vein 12 and are therefore exposed to its temperature. Two of these outgrowths 43 are wide enough to extend opposite a stage of movable blades 6 respective.
  • the ferrule 11 is therefore here formed at the same time by the protrusions 43 and 43 'and by the rings 44 supporting stationary vanes 13; the rings 44 end front and back with lips 45 which enter the grooves of the protuberances 43 and 43 '.
  • mechanical systems 46 with nesting of tenon join the rings 40 to the rings 44 concentric against mutual rotations.
  • the major difference with the downstream design is that the rings 44 are continuous on a circumference while like the rings 40. It is indeed estimated that like the heating is less important upstream, and that the temperature differences between the housing 10 and the ferrule 11 are less important also it's more simple and more advantageous to have a structure similar for both, the risks of deformation and excessive stresses being reduced.
  • the material used has a coefficient of less expansion than that used for build the downstream of the casing, because we observe that the slower expansions these materials undergo slightly regulate the evolution of dilation during the transitional phases and allow finally to better control the games at the end of the blade moving blades 6.
  • An alloy of the Inconel 909 type may be recommended or a TiAl type intermetallic.
  • the rotor 1 can be constructed in a material with a coefficient of expansion close of that used for the stator rings 40 in look, for example a titanium alloy.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP00401609A 1999-06-10 2000-06-08 Gehäuse für einen Hochdruckkompressor Expired - Lifetime EP1059420B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR9907315A FR2794816B1 (fr) 1999-06-10 1999-06-10 Stator de compresseur a haute pression
FR9907315 1999-06-10

Publications (2)

Publication Number Publication Date
EP1059420A1 true EP1059420A1 (de) 2000-12-13
EP1059420B1 EP1059420B1 (de) 2004-12-08

Family

ID=9546602

Family Applications (1)

Application Number Title Priority Date Filing Date
EP00401609A Expired - Lifetime EP1059420B1 (de) 1999-06-10 2000-06-08 Gehäuse für einen Hochdruckkompressor

Country Status (5)

Country Link
US (1) US6390771B1 (de)
EP (1) EP1059420B1 (de)
JP (1) JP4124552B2 (de)
DE (1) DE60016505T2 (de)
FR (1) FR2794816B1 (de)

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1561998A1 (de) 2004-02-05 2005-08-10 Snecma Moteurs Diffusor für Turbinentriebwerk
FR2913051A1 (fr) * 2007-02-28 2008-08-29 Snecma Sa Etage de turbine dans une turbomachine
EP2071133A1 (de) * 2007-12-14 2009-06-17 Snecma Turbotriebwerksmodul, das mit einer Vorrichtung zur Verbesserung des Radialspiels ausgestattet ist
WO2009123301A2 (en) * 2008-03-31 2009-10-08 Mitsubishi Heavy Industries, Ltd. Rotary machine
WO2010026182A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
WO2010026181A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
WO2010026179A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
CN102705254A (zh) * 2010-11-05 2012-10-03 通用电气公司 防护罩渗漏盖
WO2013162752A1 (en) 2012-04-24 2013-10-31 United Technologies Corporation Thermal management system for a gas turbine engine
FR3086323A1 (fr) 2018-09-24 2020-03-27 Safran Aircraft Engines Carter interne de turmomachine a isolation thermique amelioree

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1118806A1 (de) * 2000-01-20 2001-07-25 Siemens Aktiengesellschaft Thermisch belastbare Wand und Verfahren zur Abdichtung eines Spaltes in einer thermisch belasteten Wand
DE102004016222A1 (de) * 2004-03-26 2005-10-06 Rolls-Royce Deutschland Ltd & Co Kg Anordnung zur selbsttätigen Laufspalteinstellung bei einer zwei- oder mehrstufigen Turbine
FR2887939B1 (fr) * 2005-06-29 2016-09-30 Soc Nat D'etude Et De Construction De Moteurs D'aviation Snecma Compresseur multi-etages de turbomachine
US7604455B2 (en) * 2006-08-15 2009-10-20 Siemens Energy, Inc. Rotor disc assembly with abrasive insert
US7704038B2 (en) * 2006-11-28 2010-04-27 General Electric Company Method and apparatus to facilitate reducing losses in turbine engines
FR2925108B1 (fr) * 2007-12-14 2013-05-03 Snecma Module de turbomachine muni d'un dispositif d'amelioration des jeux radiaux
US8613593B2 (en) * 2008-12-30 2013-12-24 Rolls-Royce North American Technologies Inc. Engine case system for a gas turbine engine
JP4856257B2 (ja) * 2010-03-24 2012-01-18 川崎重工業株式会社 タービンロータのシール構造
US9091172B2 (en) 2010-12-28 2015-07-28 Rolls-Royce Corporation Rotor with cooling passage
US9115600B2 (en) * 2011-08-30 2015-08-25 Siemens Energy, Inc. Insulated wall section
US20140286766A1 (en) * 2012-09-11 2014-09-25 General Electric Company Compressor Casing Assembly Providing Access To Compressor Blade Sealing Assembly
US10539153B2 (en) * 2017-03-14 2020-01-21 General Electric Company Clipped heat shield assembly
US10767485B2 (en) * 2018-01-08 2020-09-08 Raytheon Technologies Corporation Radial cooling system for gas turbine engine compressors
US20200072070A1 (en) * 2018-09-05 2020-03-05 United Technologies Corporation Unified boas support and vane platform
US11174742B2 (en) 2019-07-19 2021-11-16 Rolls-Royce Plc Turbine section of a gas turbine engine with ceramic matrix composite vanes

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Publication number Priority date Publication date Assignee Title
US3854843A (en) * 1971-12-01 1974-12-17 R Penny Composite elongate member having a predetermined effective coefficient of linear expansion
US4101242A (en) * 1975-06-20 1978-07-18 Rolls-Royce Limited Matching thermal expansion of components of turbo-machines
US4578942A (en) * 1983-05-02 1986-04-01 Mtu Motoren-Und Turbinen-Union Muenchen Gmbh Gas turbine engine having a minimal blade tip clearance
US4805398A (en) * 1986-10-01 1989-02-21 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S. N. E. C. M. A." Turbo-machine with device for automatically controlling the rate of flow of turbine ventilation air
US5127794A (en) * 1990-09-12 1992-07-07 United Technologies Corporation Compressor case with controlled thermal environment
US5160241A (en) * 1991-09-09 1992-11-03 General Electric Company Multi-port air channeling assembly
US5314303A (en) * 1992-01-08 1994-05-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Device for checking the clearances of a gas turbine compressor casing

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DE1285255B (de) * 1964-10-28 1968-12-12 Bergmann Borsig Veb Waermebeweglich aufgehaengte Leitgittersegmente von Axialgasturbinen
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
FR2695164B1 (fr) * 1992-08-26 1994-11-04 Snecma Turbomachine munie d'un dispositif empêchant une circulation longitudinale de gaz autour des étages d'aubes de redressement.
US5653581A (en) * 1994-11-29 1997-08-05 United Technologies Corporation Case-tied joint for compressor stators
US5553999A (en) * 1995-06-06 1996-09-10 General Electric Company Sealable turbine shroud hanger
US6109868A (en) * 1998-12-07 2000-08-29 General Electric Company Reduced-length high flow interstage air extraction

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3854843A (en) * 1971-12-01 1974-12-17 R Penny Composite elongate member having a predetermined effective coefficient of linear expansion
US4101242A (en) * 1975-06-20 1978-07-18 Rolls-Royce Limited Matching thermal expansion of components of turbo-machines
US4578942A (en) * 1983-05-02 1986-04-01 Mtu Motoren-Und Turbinen-Union Muenchen Gmbh Gas turbine engine having a minimal blade tip clearance
US4805398A (en) * 1986-10-01 1989-02-21 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S. N. E. C. M. A." Turbo-machine with device for automatically controlling the rate of flow of turbine ventilation air
US5127794A (en) * 1990-09-12 1992-07-07 United Technologies Corporation Compressor case with controlled thermal environment
US5160241A (en) * 1991-09-09 1992-11-03 General Electric Company Multi-port air channeling assembly
US5314303A (en) * 1992-01-08 1994-05-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Device for checking the clearances of a gas turbine compressor casing

Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1561998A1 (de) 2004-02-05 2005-08-10 Snecma Moteurs Diffusor für Turbinentriebwerk
EP1561998B1 (de) * 2004-02-05 2012-03-07 Snecma Diffusor für Turbinentriebwerk
FR2913051A1 (fr) * 2007-02-28 2008-08-29 Snecma Sa Etage de turbine dans une turbomachine
EP1965034A1 (de) 2007-02-28 2008-09-03 Snecma Turbinenstufe in einer Strömungsmaschine
US8403636B2 (en) 2007-02-28 2013-03-26 Snecma Turbine stage in a turbomachine
EP2071133A1 (de) * 2007-12-14 2009-06-17 Snecma Turbotriebwerksmodul, das mit einer Vorrichtung zur Verbesserung des Radialspiels ausgestattet ist
FR2925109A1 (fr) * 2007-12-14 2009-06-19 Snecma Sa Module de turbomachine muni d'un dispositif d'amelioration des jeux radiaux
CN101952557A (zh) * 2008-03-31 2011-01-19 三菱重工业株式会社 回转机械
WO2009123301A2 (en) * 2008-03-31 2009-10-08 Mitsubishi Heavy Industries, Ltd. Rotary machine
WO2009123301A3 (en) * 2008-03-31 2010-09-16 Mitsubishi Heavy Industries, Ltd. Rotary machine
RU2483218C2 (ru) * 2008-03-31 2013-05-27 Мицубиси Хеви Индастрис, Лтд. Турбина
WO2010026182A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
WO2010026179A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
WO2010026181A1 (fr) * 2008-09-05 2010-03-11 Snecma Procede de fabrication d'une piece thermomecanique de revolution circulaire comportant un substrat porteur a base de titane revetu d'acier ou superalliage, carter de compresseur de turbomachine resistant au feu de titane obtenu selon ce procede
US8888448B2 (en) 2008-09-05 2014-11-18 Snecma Method for the manufacture of a circular revolution thermomechanical part including a titanium-based load-bearing substrate lined with steel or superalloy, a turbomachine compressor housing which is resistant to titanium fire obtained according to this method
CN102705254A (zh) * 2010-11-05 2012-10-03 通用电气公司 防护罩渗漏盖
CN102705254B (zh) * 2010-11-05 2016-08-31 通用电气公司 用于引导漏出空气的系统及方法
WO2013162752A1 (en) 2012-04-24 2013-10-31 United Technologies Corporation Thermal management system for a gas turbine engine
EP2841753A4 (de) * 2012-04-24 2016-10-19 United Technologies Corp Wärmemanagementsystem für einen gasturbinenmotor
FR3086323A1 (fr) 2018-09-24 2020-03-27 Safran Aircraft Engines Carter interne de turmomachine a isolation thermique amelioree
WO2020065178A1 (fr) 2018-09-24 2020-04-02 Safran Aircraft Engines Carter interne de turbomachine à isolation thermique améliorée
US11566538B2 (en) 2018-09-24 2023-01-31 Safran Aircraft Engines Internal turbomachine casing having improved thermal insulation

Also Published As

Publication number Publication date
DE60016505D1 (de) 2005-01-13
FR2794816B1 (fr) 2001-07-06
JP4124552B2 (ja) 2008-07-23
EP1059420B1 (de) 2004-12-08
US6390771B1 (en) 2002-05-21
JP2001012396A (ja) 2001-01-16
DE60016505T2 (de) 2005-11-03
FR2794816A1 (fr) 2000-12-15

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