EP1811131A2 - Anordnung von Statorsektoren für einen Verdichter eines Turbotriebwerks - Google Patents

Anordnung von Statorsektoren für einen Verdichter eines Turbotriebwerks Download PDF

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Publication number
EP1811131A2
EP1811131A2 EP07290031A EP07290031A EP1811131A2 EP 1811131 A2 EP1811131 A2 EP 1811131A2 EP 07290031 A EP07290031 A EP 07290031A EP 07290031 A EP07290031 A EP 07290031A EP 1811131 A2 EP1811131 A2 EP 1811131A2
Authority
EP
European Patent Office
Prior art keywords
flange
blades
outer shell
assembly according
annular
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
EP07290031A
Other languages
English (en)
French (fr)
Other versions
EP1811131B1 (de
EP1811131A3 (de
Inventor
Olivier Abgrall
Yvon Cloarec
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 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 SNECMA SAS filed Critical SNECMA SAS
Publication of EP1811131A2 publication Critical patent/EP1811131A2/de
Publication of EP1811131A3 publication Critical patent/EP1811131A3/de
Application granted granted Critical
Publication of EP1811131B1 publication Critical patent/EP1811131B1/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F04D29/542—Bladed diffusers
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/50—Building or constructing in particular ways
    • F05D2230/53—Building or constructing in particular ways by integrally manufacturing a component, e.g. by milling from a billet or one piece construction

Definitions

  • the present invention relates to a set of sectored fixed rectifiers for a compressor of a turbomachine, such as an airplane turbojet or turboprop.
  • a turbomachine compressor comprises a plurality of compression stages each comprising an annular row of moving blades mounted on a shaft of the turbomachine, and an annular row of fixed rectifiers carried by an outer casing.
  • Each annular row of fixed rectifiers is sectored and formed of sets of rectifiers circumferentially mounted end to end around the axis of the compressor, each set of rectifiers comprising two coaxial ferrules interconnected by radial blades, and being formed either of a single casting, either by fixing the ends of the blades on the ferrules.
  • a foundry assembly has a smaller axial footprint than an assembly formed by attaching the blades to the ferrules, but the leading and trailing edges of the blades of this assembly of a workpiece are connected to parts of the ferrule external which are themselves connected to annular mounting tabs on the outer casing and are therefore thick and very rigid.
  • the invention is intended to avoid this disadvantage while retaining the advantages of fixed sets of rectifiers formed of a casting.
  • a set of sectorized fixed rectifiers for a turbomachine compressor made in one piece and comprising two inner and outer shrouds extending coaxially one inside the other, radial blades s extending between the ferrules and connected by their radial ends to the ferrules, and two outer annular tabs carried by the outer ferrule and extending outside thereof, for mounting the set of rectifiers on a housing, characterized in that the leading or trailing edges of the blades are connected to areas of the outer shell less rigid than those connected to the annular mounting tabs.
  • the connection of the leading edges and / or leakage of the blades to parts of the outer shell which are less rigid than those connected to the annular mounting lugs makes it possible to better pass the forces between the edges of the ring. attack and leakage of the blades and the ferrule and thus to support the ferrule at least a portion of the stresses at which the leading and trailing edges of the blades are subjected in operation. This results in a significant increase in the life of these sets of fixed rectifiers.
  • At least one of the axial ends of the outer shell comprises a rim which extends substantially parallel to one of the annular mounting lugs and to which the leading edges or trailing edges of the blades.
  • the shapes and dimensions of the or each rim of the outer shell are determined so that this flange has sufficient flexibility to better distribute the constraints of the leading or trailing edges of the blades in operation.
  • the upstream end of the outer shell comprises a flange which is connected to the leading edges of the blades and which extends substantially parallel to the upstream annular lug.
  • downstream end of the shell comprises a flange which is connected to the trailing edges of the blades and which extends substantially parallel to the downstream annular tab.
  • each annular mounting lug which extends substantially parallel to such a rim of the outer shell is connected to a median portion of the shell, which may have a different thickness or radial dimension and for example greater than that of the rim.
  • the median portion of the outer shell may have a thickness or radial dimension optimized for the eigenfrequencies of the blades and ferrules while also improving the transmission of stresses between the blades and the outer shell.
  • the invention also relates to a turbomachine compressor, characterized in that it comprises at least one annular row of fixed rectifiers formed of sets of rectifiers as described above, mounted circumferentially end to end around the axis of the compressor , and a turbomachine, such as a turbojet engine or an airplane turboprop, comprising such a compressor.
  • the compressor 10 of FIG. 1 comprises several compression stages of which only two have been shown, each stage comprising an annular row of moving blades 12 whose radially inner ends are fixed on a disk 14 carried by a rotor shaft, shown, and an annular row of fixed rectifiers 16, arranged downstream of the annular row of moving blades 12 and carried by an outer cylindrical casing 18.
  • the annular rows of stationary rectifiers 16 are sectored and formed of sets of rectifiers which are mounted circumferentially end to end around the axis of the compressor.
  • Each of these sets of rectifiers comprises two internal and external coaxial ferrules 20, for example in a portion of a cylinder, which extend one inside the other and which are connected to one another by blades. 24.
  • These blades 24 have a concave inner surface or intrados and a convex or extrados outer surface which are connected to their upstream and downstream ends forming edges 26 and 28 of attack and leakage of the air flowing in the compressor.
  • Each set of fixed rectifiers is hooked on the outer casing 18 by means of two external annular tabs 30 formed at the axial ends of the outer shell 22, each annular tab 30 having an annular portion 32 which extends substantially radially towards the outside of the ferrule end 22, and a substantially cylindrical portion portion 34 which extends upstream or downstream respectively from the radially outer end of the annular portion 32 and which is engaged in a corresponding annular groove 36 of the housing.
  • the inner surface of the outer shell 22 is aligned with the inner surface of revolution of the casing 18.
  • a block of material 38 is fixed on the inner surface of the inner shell 20 and intended to cooperate in sealing with annular ribs 40 of the rotor shaft of the compressor, to prevent the passage of gas between the inner ring 20 and the rotor shaft.
  • the set of rectifiers of FIG. 1 is formed of a single piece, in particular of casting, which makes it possible to minimize the axial dimension of the outer shell 22 and thus the axial space requirement of the rectifier assembly by bringing the legs external 30 as close to the leading edges 26 and leakage 28 of the radial blades.
  • the leading edges 26 and trailing edges 28 of the blades are thus connected to thick and rigid parts of the shell which are not flexible enough to partially absorb the stresses to which the leading and trailing edges of the blades are subjected. operation.
  • the set of rectifiers of FIG. 2 is formed by assembling radial blades 24 on ferrules 20, 22, more precisely by fitting and welding or brazing the ends of the blades 24 in corresponding orifices of the ferrules 20, 22.
  • the bulk axial of such an assembly is greater than that of the assembly of Figure 1 because the attachment tabs 30 of the assembly are necessarily spaced upstream and downstream respectively, the mounting holes of the blades radials.
  • this embodiment allows the leading and trailing edges of the blades to be connected to relatively thin ferrule portions which are sufficiently flexible to absorb a portion of the stresses to which the leading edges 26 and 40 are subjected. leakage 28 blades in operation.
  • the present invention makes it possible to combine the advantages and to avoid the disadvantages of these two embodiments.
  • the outer ring 52 of the set of rectifiers 50 comprises a downstream flange 54 substantially in a portion of a cylinder which is aligned axially with the remainder of the shell and which extends substantially parallel to the downstream annular tab 56 and inside thereof.
  • This tab 56 comprises an annular portion 58 which extends substantially radially outwardly from the shell 52, and a portion 60 substantially in a portion of a cylinder or cone which extends downstream. from the radially outer end of the annular portion 58 and which is intended to be engaged in an annular groove of the housing 18.
  • the radial portion 58 of the downstream flap 56 is connected to the outer shell 52 upstream of the trailing edges 28 of the radial blades 24, which are connected to the rim 54 forming the downstream end of the outer shell 52.
  • its thickness may be reduced to give it a certain flexibility, which allows it to absorb some of the constraints applied to the trailing edges of the blades 24 in operation.
  • the thickness of the flange 54 may be substantially equal to that of the remainder of the ferrule (outside the connection zones of the external fastening tabs) or lower.
  • the portion 60 of the downstream leg 56 also comprises a radially outer annular rib 62 intended to abut on a corresponding surface of the housing when the portion 60 of the tab is engaged in the groove of the housing. In the view of FIG. 1, this surface is radial and formed by a cylindrical rim of the casing 18.
  • the inner ferrule 20, the blades 24 and the upstream annular flange 30 are similar to those of the assembly of FIG. 1.
  • the assembly 50 shown in FIG. 4 differs from that of FIG. 3 in that the outer shell 52 also comprises an upstream rim 64 that is substantially in the form of a cylinder portion that is aligned axially with the remainder of the shell and that extends substantially in parallel to the upstream annular tab 66 for hooking the assembly and inside thereof.
  • This lug 66 has an annular portion 68 that extends substantially radially outwardly from a portion of the shell located downstream of the zone of connection of the leading edges 26 of the blades, and a second portion 70 substantially in portion of cylinder which extends substantially axially upstream from the radially outer end of the annular portion 68 and which is intended to be engaged in an annular groove of the housing 18.
  • the upstream and downstream annular tabs 66 and 56 are separated from each other by a median portion of outer shell 52 which has substantially the same thickness or radial dimension as the upstream and downstream edges 64 and 54 of this ferrule. This makes it possible to optimize the eigenfrequencies of the blades 24 and the outer shell 52 while improving the transmission of stresses between the leading and trailing edges of the blades and the outer shell.
  • the upstream annular and downstream annular tabs 56 are connected to a median part 72 of the outer shell which has a thickness or radial dimension that is clearly greater than that of the upstream and downstream flanges. of the outer shell.
  • the first radial portions 58, 68 of the upstream and downstream lugs 66, respectively, are formed by the thick medial portion of the outer shell 52. This embodiment also makes it possible to improve the transmission of stresses between the blades and the outer shell.
  • the invention is not limited to the embodiments that have been described in the foregoing and shown in the accompanying drawings.
  • the set of rectifiers of Figure 3 could be formed with a flange 54 at its upstream end and with an annular tab 30 at its downstream end, as opposed to what is shown.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP07290031.9A 2006-01-24 2007-01-10 Anordnung von Statorsektoren für einen Verdichter eines Turbotriebwerks Active EP1811131B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0600616A FR2896548B1 (fr) 2006-01-24 2006-01-24 Ensemble de redresseurs fixes sectorise pour un compresseur de turbomachine

Publications (3)

Publication Number Publication Date
EP1811131A2 true EP1811131A2 (de) 2007-07-25
EP1811131A3 EP1811131A3 (de) 2008-09-24
EP1811131B1 EP1811131B1 (de) 2017-05-17

Family

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Family Applications (1)

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EP07290031.9A Active EP1811131B1 (de) 2006-01-24 2007-01-10 Anordnung von Statorsektoren für einen Verdichter eines Turbotriebwerks

Country Status (4)

Country Link
US (1) US7946811B2 (de)
EP (1) EP1811131B1 (de)
FR (1) FR2896548B1 (de)
RU (1) RU2439338C2 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2011157957A1 (fr) 2010-06-18 2011-12-22 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
WO2011157956A1 (fr) 2010-06-18 2011-12-22 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
FR2984428A1 (fr) * 2011-12-19 2013-06-21 Snecma Redresseur de compresseur pour turbomachine.
WO2016142631A1 (fr) * 2015-03-11 2016-09-15 Microturbo Réalisation de demi-etages de redresseurs monoblocs, par fabrication additive
FR3048015A1 (fr) * 2016-02-19 2017-08-25 Snecma Aube de turbomachine, comprenant un pied aux concentrations de contrainte reduites
WO2020157405A1 (fr) * 2019-01-30 2020-08-06 Safran Aircraft Engines Secteur de stator de turbomachine a zones soumises a des contraintes élevées assouplies

Families Citing this family (8)

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Publication number Priority date Publication date Assignee Title
WO2009115384A1 (de) * 2008-03-19 2009-09-24 Alstom Technology Ltd Leitschaufel mit hakenförmigem befestigungselement für eine gasturbine
FR2985792B1 (fr) 2012-01-18 2014-02-07 Snecma Secteur angulaire de redresseur a amortissement de vibrations par coin pour compresseur de turbomachine
US20130333350A1 (en) * 2012-06-19 2013-12-19 Nicholas D. Stilin Airfoil including adhesively bonded shroud
US11035238B2 (en) 2012-06-19 2021-06-15 Raytheon Technologies Corporation Airfoil including adhesively bonded shroud
US9702252B2 (en) 2012-12-19 2017-07-11 Honeywell International Inc. Turbine nozzles with slip joints and methods for the production thereof
GB2556054A (en) * 2016-11-16 2018-05-23 Rolls Royce Plc Compressor stage
US20180340438A1 (en) * 2017-05-01 2018-11-29 General Electric Company Turbine Nozzle-To-Shroud Interface
CN110030037B (zh) * 2018-01-11 2021-08-13 中国航发商用航空发动机有限责任公司 涡轮导向叶片、涡轮导向叶片组件以及核心机

Citations (5)

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Publication number Priority date Publication date Assignee Title
FR1252179A (fr) * 1959-12-17 1961-01-27 Snecma Perfectionnement aux stators de machines à fluide à écoulement axial
FR2664944A1 (fr) * 1990-07-18 1992-01-24 Snecma Compresseur forme notamment de redresseurs en couronne et procede de montage de ce compresseur.
FR2674909A1 (fr) * 1991-04-03 1992-10-09 Snecma Stator de compresseur de turbomachine a aubes demontables.
EP1505259A1 (de) * 2003-08-08 2005-02-09 ROLLS-ROYCE plc Vorrichtung zur Befestigung einer nicht-rotierenden Komponente einer Gasturbine
US20050111969A1 (en) * 2003-11-20 2005-05-26 General Electric Company Apparatus and methods for removing and installing a selected nozzle segment of a gas turbine in an axial direction

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US3365173A (en) * 1966-02-28 1968-01-23 Gen Electric Stator structure
US4856963A (en) * 1988-03-23 1989-08-15 United Technologies Corporation Stator assembly for an axial flow rotary machine
US6514041B1 (en) * 2001-09-12 2003-02-04 Alstom (Switzerland) Ltd Carrier for guide vane and heat shield segment
US6910854B2 (en) * 2002-10-08 2005-06-28 United Technologies Corporation Leak resistant vane cluster
US7040857B2 (en) * 2004-04-14 2006-05-09 General Electric Company Flexible seal assembly between gas turbine components and methods of installation

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1252179A (fr) * 1959-12-17 1961-01-27 Snecma Perfectionnement aux stators de machines à fluide à écoulement axial
FR2664944A1 (fr) * 1990-07-18 1992-01-24 Snecma Compresseur forme notamment de redresseurs en couronne et procede de montage de ce compresseur.
FR2674909A1 (fr) * 1991-04-03 1992-10-09 Snecma Stator de compresseur de turbomachine a aubes demontables.
EP1505259A1 (de) * 2003-08-08 2005-02-09 ROLLS-ROYCE plc Vorrichtung zur Befestigung einer nicht-rotierenden Komponente einer Gasturbine
US20050111969A1 (en) * 2003-11-20 2005-05-26 General Electric Company Apparatus and methods for removing and installing a selected nozzle segment of a gas turbine in an axial direction

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105134304B (zh) * 2010-06-18 2017-04-12 斯奈克玛 用于涡轮引擎压缩机、涡轮引擎定子的定子的角区部以及包括此区部的涡轮引擎
CN105134304A (zh) * 2010-06-18 2015-12-09 斯奈克玛 用于涡轮引擎压缩机、涡轮引擎定子的定子的角区部以及包括此区部的涡轮引擎
US9222363B2 (en) 2010-06-18 2015-12-29 Snecma Angular sector of a stator for a turbine engine compressor, a turbine engine stator, and a turbine engine including such a sector
FR2961554A1 (fr) * 2010-06-18 2011-12-23 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
CN103038454A (zh) * 2010-06-18 2013-04-10 斯奈克玛 一种涡轮发动机压缩机定子的角扇形片,包含该角扇形片的涡轮发动机定子和涡轮发动机
US9228449B2 (en) 2010-06-18 2016-01-05 Snecma Angular sector of a stator for a turbine engine compressor, a turbine engine stator, and a turbine engine including such a sector
WO2011157956A1 (fr) 2010-06-18 2011-12-22 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
CN103038454B (zh) * 2010-06-18 2014-12-31 斯奈克玛 一种涡轮发动机压缩机定子的角扇形片及包含该角扇形片的涡轮发动机定子和涡轮发动机
EP2949869A1 (de) 2010-06-18 2015-12-02 Snecma Leitschaufelsektor für verdichter einer turbomaschine, statoranordnung einer turbomaschine und turbomaschine umfassend eine solche statoranordnung
RU2584078C2 (ru) * 2010-06-18 2016-05-20 Снекма Угловой сектор статора для компрессора газотурбинного двигателя, статор газотурбинного двигателя и газотурбинный двигатель, включающий в себя такой сектор
FR2961553A1 (fr) * 2010-06-18 2011-12-23 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
WO2011157957A1 (fr) 2010-06-18 2011-12-22 Snecma Secteur angulaire de redresseur pour compresseur de turbomachine, redresseur de turbomachine et turbomachine comprenant un tel secteur
US9702259B2 (en) 2011-12-19 2017-07-11 Snecma Turbomachine compressor guide vanes assembly
WO2013093337A1 (fr) * 2011-12-19 2013-06-27 Snecma Redresseur de compresseur pour turbomachine
FR2984428A1 (fr) * 2011-12-19 2013-06-21 Snecma Redresseur de compresseur pour turbomachine.
WO2016142631A1 (fr) * 2015-03-11 2016-09-15 Microturbo Réalisation de demi-etages de redresseurs monoblocs, par fabrication additive
FR3033602A1 (fr) * 2015-03-11 2016-09-16 Microturbo Realisation d'etages de redresseurs semi-monoblocs, par fabrication additive
FR3048015A1 (fr) * 2016-02-19 2017-08-25 Snecma Aube de turbomachine, comprenant un pied aux concentrations de contrainte reduites
US10858957B2 (en) 2016-02-19 2020-12-08 Safran Aircraft Engines Turbomachine blade, comprising a root with reduced stress concentrations
CN113366192B (zh) * 2019-01-30 2024-04-02 赛峰航空器发动机 具有承受高应力的柔性区域的涡轮机定子扇区
WO2020157405A1 (fr) * 2019-01-30 2020-08-06 Safran Aircraft Engines Secteur de stator de turbomachine a zones soumises a des contraintes élevées assouplies
CN113366192A (zh) * 2019-01-30 2021-09-07 赛峰航空器发动机 具有承受高应力的柔性区域的涡轮机定子扇区
US11767767B2 (en) 2019-01-30 2023-09-26 Safran Aircraft Engines Turbomachine stator sector having flexible regions subjected to high stress

Also Published As

Publication number Publication date
US20070172349A1 (en) 2007-07-26
EP1811131B1 (de) 2017-05-17
US7946811B2 (en) 2011-05-24
FR2896548A1 (fr) 2007-07-27
RU2439338C2 (ru) 2012-01-10
FR2896548B1 (fr) 2011-05-27
EP1811131A3 (de) 2008-09-24
RU2007102521A (ru) 2008-07-27

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