EP1759090A1 - Aube comportant une zone de transition - Google Patents

Aube comportant une zone de transition

Info

Publication number
EP1759090A1
EP1759090A1 EP05749220A EP05749220A EP1759090A1 EP 1759090 A1 EP1759090 A1 EP 1759090A1 EP 05749220 A EP05749220 A EP 05749220A EP 05749220 A EP05749220 A EP 05749220A EP 1759090 A1 EP1759090 A1 EP 1759090A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
thickening
cut
end region
radius
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Withdrawn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
EP05749220A
Other languages
German (de)
English (en)
Inventor
Martin Hoeger
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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 MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Publication of EP1759090A1 publication Critical patent/EP1759090A1/fr
Withdrawn legal-status Critical Current

Links

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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • 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 invention relates to an airfoil of a guide blade or moving blade of a turbomachine according to the preamble of claim 1. Furthermore, the invention relates to a turbomachine according to the preamble of claim 13.
  • EP 0 798 447 B1 discloses an airfoil for turbomachines, in particular for gas turbines, which extend in a radial direction in an annular duct or flow duct, the aerofoils having a thickening of the aerofoil profile on at least one delimiting side wall of the annular duct. According to EP 0 798 447 B1, this thickening is formed by an increased leading edge nose radius and / or an increased leading edge wedge angle and / or an increased trailing edge wedge angle and / or a greater absolute profile thickness of profile cuts adjoining the limiting side wall compared to a reference profile cut.
  • the present invention is based on the problem of creating a novel airfoil of a guide blade or rotor blade of a turbomachine.
  • the thickening is cut in the area of a front edge of the airfoil to form a mirror surface or base surface.
  • the thickening in the area of the front edge is cut with a straight or flat cut in such a way that the cutting direction of the cut runs in the circumferential direction or transverse to the axial direction of the turbomachine.
  • the thickening in the area of the front edge is cut with a circular segment or cylinder segment shaped cut such that a tangent applied to the foremost part of the cut runs in the circumferential direction or transverse to the axial direction of the turbomachine.
  • the or each thickening is preferably formed by a large radius of curvature at the radially inner end region and / or radially outer end region of the airfoil, the ratio of the radius of curvature to a chord length of the airfoil being between 2% and 10%.
  • turbomachine according to the invention is defined in independent claim 13.
  • FIG. 1 shows an airfoil according to the prior art in a highly schematic, perspective side view
  • FIG. 2 shows the airfoil of FIG. 1 in viewing direction II of FIG. 1;
  • FIG. 3 shows an airfoil according to the prior art in a representation analogous to FIG. 2;
  • FIG. 4 shows a section through the airfoil of FIG. 3 in the cutting direction III-III;
  • FIG. 5 shows an airfoil according to the invention in a highly schematic, perspective side view
  • FIG. 6 shows the airfoil of FIG. 5 in viewing direction VI of FIG. 5;
  • FIG. 7 shows a second airfoil according to the invention in a representation analogous to FIG. 6;
  • FIGS. 8 shows a third airfoil according to the invention in a representation analogous to FIGS. 6 and 7.
  • Fig. 1 shows a known from the prior art blade 10 of a blade in a highly schematic representation
  • the blade 10 of the blade according to FIG. 1 has a flow-mechanically advantageous thickening 11 of the blade profile at a radially inner, hub-side end region.
  • the thickening 11 extends according to FIG 3 are the chord length 1 of the airfoil, which is defined by the distance between the front edge 12 and the rear edge 13, and the maximum profile thickness d and others x of the airfoil.
  • the thickening 11 at the hub-side end of the airfoil 10 of the rotor blade which is not shown further, is achieved in that a hub-side profile cut of the airfoil 10 has a larger absolute profile thickness d than a reference profile cut, the contour of the thickening 11 is defined according to FIG. 4 by a radius of curvature r. With such a thickening 11, secondary flow losses can be minimized.
  • the thickening 11 shown in FIGS. 1 to 4 merely represents a possible variant of thickening that minimizes secondary flow. Further thickening that minimizes secondary flow can result from an enlargement of the profile parameters leading edge nose radius and / or leading edge wedge angle and / or trailing edge wedge angle and / or absolute profile thickness. With regard to the details of such thickenings, reference is made to EP 0 798 447 B1, the disclosure content of which is explicitly referred to here by reference.
  • the thickening 11 in the region of the front edge 12 of the airfoil 10 is cut with a straight or flat cut such that the cutting direction of the cut runs in the circumferential direction (arrow 15) of the turbomachine .
  • the flat mirror surface 14 that is set in this case runs parallel to the circumferential direction 15.
  • the cutting direction for providing the flat mirror surface 14 is transverse or perpendicular to the axial direction (Arrow 16) of the turbomachine.
  • the thickening 11 is in the embodiment of FIGS. 5 and 6, as described in connection with FIGS. 1 to 4, continuously formed around the airfoil and characterized by a large radius of curvature r at the radially inner end region or radially outer end region of the airfoil, wherein
  • the radius of curvature r is dimensioned such that the ratio of the radius of curvature r to the chord length 1 of the airfoil 10 is between 2% and 10%, preferably between 4% and 8%, particularly preferably between 5% and 7%.
  • the ratio of rounding radius r to chord length is preferably 1 6%.
  • FIG. 7 shows a further exemplary embodiment of an airfoil 10 according to the invention, wherein in the exemplary embodiment of FIG. 7 the airfoil 10 also has a thickening 11 on the radially inner, end region on the hub side and / or on the radially outer, end region on the housing side, the thickening 11 in turn Area of the front edge 12 is cut to form a mirror surface 17.
  • the thickening 11 in the region of the front edge 12 is not in the exemplary embodiment in FIG. 7 cut through a flat cut, but rather through a circular segment or cylinder segment shaped cut.
  • the thickening 11 is cut through the circular segment or cylindrical segment-shaped cut in such a way that a tangent 18 applied to the foremost part of the cut or the mirror surface 17 that is formed parallel to the circumferential direction (arrow 15) and transversely or perpendicularly to the axial direction (arrow 16 ) of the turbomachine.
  • FIG. 8 shows an airfoil 10 designed according to the invention with a variable radius of curvature r, which decreases starting from the front edge 12 in the direction of the rear edge 13 of the airfoil.
  • the airfoil 10 is in turn cut off at the front edge 12 to form a mirror surface or base surface, a cut in the form of a segment of a circle or a cylinder segment being formed in the exemplary embodiment in FIG. 8 as well as in the exemplary embodiment in FIG. 7 Formation of the mirror surface 17 is executed.
  • a variable radius of curvature r in the region of the thickening 11, secondary flow loss can be minimized again.
  • the airfoils according to the invention are preferably used in guide vanes or moving blades of a turbine or a compressor of a gas turbine, in particular of a gas turbine aircraft engine.
  • the blades of rotating rotor blades have thickenings designed according to the invention only in the region of the hub-side end, fixed guide blades can have such a thickening both at the radially inner end on the hub side and at the radially outer end on the housing side.
  • the thickenings in the area of the front edge are cut to form a mirror surface or base surface.
  • a large radius of curvature is used in the sense of the invention.
  • secondary flow losses can be minimized in this way, and on the other hand, the axial overall depth or overall length of the turbomachine equipped with such blades or blades is not increased.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Control Of Turbines (AREA)

Abstract

L'invention concerne une aube d'un ensemble aube directrice ou d'un ensemble aube mobile d'une turbomachine, en particulier d'une turbine à gaz. Le profilé de l'aube comporte au moins une partie plus épaisse, avantageuse en termes de mécanique des fluides, au niveau d'une zone terminale située radialement à l'intérieur et/ou d'une zone terminale située radialement à l'extérieur de l'aube. Selon l'invention, cette partie plus épaisse (11) est coupée dans la zone d'une arête antérieure (12) de l'aube, de manière à former une surface miroir (14, 17) ou une surface de base.
EP05749220A 2004-05-29 2005-04-29 Aube comportant une zone de transition Withdrawn EP1759090A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102004026386A DE102004026386A1 (de) 2004-05-29 2004-05-29 Schaufelblatt einer Strömungsmaschine sowie Strömungsmaschine
PCT/DE2005/000788 WO2005116404A1 (fr) 2004-05-29 2005-04-29 Aube comportant une zone de transition

Publications (1)

Publication Number Publication Date
EP1759090A1 true EP1759090A1 (fr) 2007-03-07

Family

ID=34969634

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05749220A Withdrawn EP1759090A1 (fr) 2004-05-29 2005-04-29 Aube comportant une zone de transition

Country Status (6)

Country Link
US (1) US20070177979A1 (fr)
EP (1) EP1759090A1 (fr)
DE (1) DE102004026386A1 (fr)
RU (1) RU2383748C2 (fr)
UA (1) UA91191C2 (fr)
WO (1) WO2005116404A1 (fr)

Families Citing this family (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2899270A1 (fr) * 2006-03-30 2007-10-05 Snecma Sa Aube de redresseur a amenagement de forme localise, secteur de redresseurs, etage de compression, compresseur et turbomachine comportant une telle aube
FR2899269A1 (fr) * 2006-03-30 2007-10-05 Snecma Sa Aube de redresseur optimisee, secteur de redresseurs, etage de compression, compresseur et turbomachine comportant une telle aube
DE102008017624A1 (de) * 2008-04-04 2009-10-08 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur aerodynamischen Ausformung der Vorderkante von Bliskschaufeln
DE102009036406A1 (de) 2009-08-06 2011-02-10 Mtu Aero Engines Gmbh Schaufelblatt
DE102010004854A1 (de) 2010-01-16 2011-07-21 MTU Aero Engines GmbH, 80995 Laufschaufel für eine Strömungsmaschine und Strömungsmaschine
US8944774B2 (en) * 2012-01-03 2015-02-03 General Electric Company Gas turbine nozzle with a flow fence
EP2811115A1 (fr) * 2013-06-05 2014-12-10 Alstom Technology Ltd Profil d'aube pour turbine à gaz, pale et aube de guidage
DE102013213416B4 (de) * 2013-07-09 2017-11-09 MTU Aero Engines AG Schaufel für eine Gasturbomaschine
JP5916826B2 (ja) * 2014-09-24 2016-05-11 三菱日立パワーシステムズ株式会社 回転機械の翼体、及びガスタービン
US9995166B2 (en) * 2014-11-21 2018-06-12 General Electric Company Turbomachine including a vane and method of assembling such turbomachine
US9890641B2 (en) 2015-01-15 2018-02-13 United Technologies Corporation Gas turbine engine truncated airfoil fillet
DE102015224283A1 (de) 2015-12-04 2017-06-08 MTU Aero Engines AG Leitschaufelcluster für eine Strömungsmaschine
US11098591B1 (en) 2019-02-04 2021-08-24 Raytheon Technologies Corporation Turbine blade with contoured fillet
DE102021123281A1 (de) 2021-09-08 2023-03-09 MTU Aero Engines AG Schaufelblatt für einen Verdichter einer Strömungsmaschine

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CH342326A (de) * 1955-03-31 1959-11-15 Kress Herwig Ing Dr Im Schaufelträger befestigte Schaufel für axial durchströmte Turbomaschinen und Verfahren zu deren Herstellung
DE3514122A1 (de) * 1985-04-19 1986-10-23 MAN Gutehoffnungshütte GmbH, 4200 Oberhausen Verfahren zur herstellung einer leitschaufel fuer ein turbinen- oder verdichter-leitrad und nach dem verfahren hergestellte leitschaufel
US6511294B1 (en) * 1999-09-23 2003-01-28 General Electric Company Reduced-stress compressor blisk flowpath
US6524070B1 (en) * 2000-08-21 2003-02-25 General Electric Company Method and apparatus for reducing rotor assembly circumferential rim stress
US6851924B2 (en) * 2002-09-27 2005-02-08 Siemens Westinghouse Power Corporation Crack-resistance vane segment member

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO2005116404A1 *

Also Published As

Publication number Publication date
US20070177979A1 (en) 2007-08-02
RU2383748C2 (ru) 2010-03-10
WO2005116404A1 (fr) 2005-12-08
UA91191C2 (ru) 2010-07-12
RU2006145067A (ru) 2008-07-10
DE102004026386A1 (de) 2005-12-22

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