WO2005106207A1 - Aube de compresseur et compresseur - Google Patents

Aube de compresseur et compresseur Download PDF

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Publication number
WO2005106207A1
WO2005106207A1 PCT/EP2005/051534 EP2005051534W WO2005106207A1 WO 2005106207 A1 WO2005106207 A1 WO 2005106207A1 EP 2005051534 W EP2005051534 W EP 2005051534W WO 2005106207 A1 WO2005106207 A1 WO 2005106207A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
profile
compressor
tip
platform
Prior art date
Application number
PCT/EP2005/051534
Other languages
German (de)
English (en)
Inventor
Reinhard MÖNIG
Original Assignee
Siemens Aktiengesellschaft
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 Siemens Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Publication of WO2005106207A1 publication Critical patent/WO2005106207A1/fr

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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
    • 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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • 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
    • 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

Definitions

  • the invention relates to a compressor blade with a blade root, a platform and a free-standing blade profile with a profile tip, according to the preamble of claim 1 and a compressor according to claim 7.
  • Free-standing guide vanes and rotor blades of an axial compressor are known.
  • the compressor blades are arranged in rings in the compressor and each have a blade profile which is arranged in the flow channel of the compressor.
  • each free end of the blade profile has a profile tip, on the end face of which a sealing lip is formed, which extends from a leading edge to a trailing edge of the blade profile.
  • the sealing lip serves to reduce a radial gap which is formed between the profile tip of the compressor blade and a component opposite this profile tip and through which lossy leakage of the flow medium of the compressor takes place.
  • Compressor e.g. a guide ring, which delimits the flow channel radially on the outside.
  • the rotor is the component which lies opposite the profile tip of a guide vane and which delimits the radial gap and thus the flow channel radially on the inside.
  • JP 60-184905 discloses a moving blade with a projecting collar which is arranged on the pressure side of the blade on the head side.
  • EP 1 225 303 A2 describes a turbine rotor blade with a blade tip that is enlarged in the axial direction Known to reduce the pressure drop on the turbine side. Because the rotor blade tip is substantially enlarged in the axial direction, the adjacent guide blades have a profile taper in the region near the platform, that is to say radially on the outside, which is proportional to the widening.
  • JP 2002-48095 shows a turbine blade with a profile cross section enlarged in the profile region near the tip, compared to its central profile region.
  • the enlargement particularly affects the leading edge.
  • the first-mentioned object is achieved by a compressor blade according to claim 1 and the object directed to the compressor is achieved by the features of claim 7.
  • the solution to the first-mentioned problem suggests that in the longitudinal section through the free-standing, i.e. Compressor blades without a shroud
  • the invention proposes a blade profile that widens in the direction of the profile tip between the suction side surface and the pressure side surface. This causes obtuse angles between the pressure side surface or the suction side surface and a surface of a flow channel boundary that is essentially perpendicular to the blade axis.
  • the Flow channel limitation can be formed by a guide ring or a circumferential surface of a rotor.
  • the side surfaces of the tip-close profile areas which are inclined also lead to a reduction in flow separation in the area of the blade inter-edge in the tip-close profile area, which consequently leads to a considerable aerodynamic improvement and thus to reduced flow losses.
  • the thickness of the leading edge and the trailing edge remain essentially unchanged in the profile area close to the point compared to the middle and platform-close profile areas.
  • the compressor blade is particularly advantageously provided with a continuous widening. This course, which is not graded in longitudinal section, requires particularly good flow behavior and low disturbances in the flow medium.
  • the expansion is expediently arranged on both sides, starting from the central profile region.
  • the profile tip has an end face in which at least one recess is provided, which is elongated between the front edge of the blade and the rear edge of the blade.
  • the blade side walls, which form the side surfaces, thus form sealing lips on the profile tip, which guide the leakage flow over the Reduce the profile tip.
  • the leakage flow in the recess which cannot be completely avoided, causes swirling, which in the manner of a labyrinth seal produces a sealing effect and thus limits the leakage flow to a smaller extent.
  • the blade profile have a profile height in the direction of the blade axis between the platform surface and the profile tip, and that the profile regions close to the tip are spaced from the platform by at least 85% of the profile height.
  • the processing takes place only in the outermost area of the profile tip.
  • the present invention is suitable for both compressor blades designed as a rotor blade and for a guide blade.
  • a compressor according to the invention has compressor blades according to one of claims 1 to 4.
  • FIG. 1 is a perspective view of a compressor blade as a moving blade
  • Fig. 2 is a profile tip of a Laufschaufei and one of the profile tip opposite housing inner wall in longitudinal section and
  • Fig. 3 shows a profile tip of a guide vane and a rotor section opposite the profile tip in longitudinal section.
  • FIG. 1 shows a compressor blade 1 as a rotor blade 3.
  • the rotor blade 3 has a blade axis 5 which corresponds to the radial direction of a rotor 6 or a rotor shaft of the compressor (not shown).
  • the blade root 7 is followed by the platform 9.
  • a blade profile 13 is arranged on the surface 11 of the platform 9 and has a profile tip 15 at its end facing away from the platform 9.
  • the blade profile 13 comprises a blade leading edge 17 and a blade trailing edge 19, between which a pressure-side blade side wall 20 and a suction-side blade side wall 22 opposite the pressure-side 'blade side wall 20 extend.
  • the surfaces of the blade side walls 20, 22 which are exposed to the flow medium form a pressure side surface 21 and a suction side surface 23.
  • the blade profile 13 along the blade axis 5 has a profile height 25 which extends between the surface 11 of the platform 9 and the profile tip 15. Along the blade axis 5, the blade profile 13 continuously merges from profile areas 27 near the platform via central profile areas 29 into profile areas 31 close to the tip. The distance of the
  • Blade leading edge 17 to blade trailing edge 19 is shown as blade width 26.
  • the intersection lines 32 of the suction side surface 23 and the pressure side surface 21 with an imaginary plane perpendicular to the blade axis 5 enclose a profile cross section 33 which is shown in dashed lines in the central profile region 29 and in the profile region 27 near the platform
  • the profile cross-sections 33 are approximately identical in shape, size and curvature for the middle and for the platform-near profile area 27, 29.
  • the profile cross section increases continuously and not in stages in comparison to that of the middle and platform-close profile regions.
  • the distance A of the cutting lines 32 which forms the imaginary plane with the pressure side surface 21 and the suction side surface 23, increases, so that a widening of the profile cross section 34 in the near-pointed direction Profile area 31 is present.
  • the profile tip 15 has a profile cross section 35 which is enlarged compared to the profile cross sections 33, 34.
  • the blade profile 13 has no or only a slight widening in the profile area 31 near the tip in the area of the blade leading edge 17 (compared to the middle or platform-like profile areas 27, 29), which increases in the direction of flow of the flow medium up to a center 28 of the blade width 26 and in further course to the blade trailing edge 19 decreases again.
  • the trailing edge of the blade 19 of the profile region 31 close to the tip is also virtually unchanged in longitudinal section compared to the trailing edge 19 of the middle or the profile region 27, 29 near the platform.
  • the blade width 26 is provided with two elongate recesses 39, which are separated from one another by a partition 41 lying between them.
  • Fig. 2 shows partially one arranged on a rotor
  • Blade 3 of a compressor with a housing inner wall 43 in a longitudinal section.
  • the rotor blade 3 is only shown with its central profile region 29 and the profile region 31 close to the point.
  • the pressure side surface 21 and the suction side surface 23 widen the blade profile 13 in
  • the blade profile 13 bifurcates into a pressure-side sealing lip 36 and a suction-side sealing lip 38 in the profile region 31 near the tip in the direction of the profile tip 15.
  • Profile area 31 with the housing inner wall 43 makes an obtuse angle ⁇ and the suction side surface 23 in the profile section 31 close to the point with the housing inner wall 43 makes an obtuse angle ⁇ .
  • Pressure ratios are represented by plus and minus in Fig. 2.
  • a leakage flow 47 is caused by the different pressures, which flows through the radial gap 45 streams.
  • the leakage flow 47 creates swirls in the recesses 39 which limit the flow velocity of the leakage flow 47 and thus reduce the leakage.
  • the rotating rotor moves the rotor blade 3 in the direction R, so that in the profile region 31 near the tip, a radially acting component 49 pushes the flow medium to be compressed away from the radial gap 45 by the obliquely inclined pressure side surface 21. This further reduces the leakage flow 47.
  • the obtuse angle CC, ⁇ results in an aerodynamic improvement of the compressor blade 1, so that the flow in the area of the blade trailing edge 19 in the tip-close profile region 31 is less likely to detach or tear off.
  • FIG. 3 shows, like FIG. 2, a compressor blade 1 with an opposite radial flow limitation.
  • a guide blade 4 is shown in this case and, instead of the housing inner wall 43, the rotor 6 with its circumferential surface 8 is partially shown.
  • a radial component 49 is generated in the flow medium flowing around the blade profile 13, so that it is led away from the radial gap 45 and the leakage flow 47 is further reduced.
  • the radial component 4 is more pronounced in the area of the pressure side surface 21 than on the suction side surface 23.
  • a compressor blade 1 in the profile region 31 close to the tip can also be designed such that only on the pressure side (or on the suction side) does the corresponding side surface 21 (23) cause the blade profile 13 to widen, so that the opposite suction side surface 23 (or the pressure side surface) 21) strikes the surface delimiting the flow channel 10 at a perpendicular angle.
  • compressor blade 1 and the compressor are suitable for use in a gas turbine.
  • the invention can also be applied to free-standing blades of a turbine.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne une aube de compresseur (1) présentant une emplanture (7), une plate-forme (9) et un profilé saillant (1) pourvu d'une extrémité (15), ces éléments étant placés de manière à se succéder le long d'un axe d'aube (5). Le profilé d'aube (13) présente une face côté admission (23) courbée ainsi qu'une face côté refoulement (21) également courbée, opposée à la face côté admission (23). Le profilé d'aube (13) présente des zones (27, 29, 31) qui sont respectivement proches de la plate-forme, centrales et proches de l'extrémité et qui se rejoignent de manière continue en direction de l'axe d'aube (5). L'objectif de l'invention est de créer une aube de compresseur qui présente un rendement particulièrement bon. A cet effet, la distance (A), en coupe longitudinale à travers l'aube de compresseur (1), entre la face côté refoulement (21) et la face côté admission (23), dans les zones de profilé (31) proches de l'extrémité augmente dans la direction de l'extrémité de profilé (15).
PCT/EP2005/051534 2004-04-27 2005-04-07 Aube de compresseur et compresseur WO2005106207A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP04010000A EP1591624A1 (fr) 2004-04-27 2004-04-27 Aube de compresseur et compresseur
EP04010000.0 2004-04-27

Publications (1)

Publication Number Publication Date
WO2005106207A1 true WO2005106207A1 (fr) 2005-11-10

Family

ID=34924757

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2005/051534 WO2005106207A1 (fr) 2004-04-27 2005-04-07 Aube de compresseur et compresseur

Country Status (2)

Country Link
EP (1) EP1591624A1 (fr)
WO (1) WO2005106207A1 (fr)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102009036406A1 (de) * 2009-08-06 2011-02-10 Mtu Aero Engines Gmbh Schaufelblatt
US8632311B2 (en) 2006-08-21 2014-01-21 General Electric Company Flared tip turbine blade
DE102012021400A1 (de) * 2012-10-31 2014-04-30 Rolls-Royce Deutschland Ltd & Co Kg Turbinenrotorschaufel einer Gasturbine
US9593584B2 (en) 2012-10-26 2017-03-14 Rolls-Royce Plc Turbine rotor blade of a gas turbine
US10107108B2 (en) 2015-04-29 2018-10-23 General Electric Company Rotor blade having a flared tip
US10267330B2 (en) 2014-06-25 2019-04-23 Siemens Aktiengesellschaft Compressor aerofoil and corresponding compressor rotor assembly

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7607893B2 (en) * 2006-08-21 2009-10-27 General Electric Company Counter tip baffle airfoil
US7686578B2 (en) * 2006-08-21 2010-03-30 General Electric Company Conformal tip baffle airfoil
FR2907157A1 (fr) * 2006-10-13 2008-04-18 Snecma Sa Aube mobile de turbomachine
GB0724612D0 (en) 2007-12-19 2008-01-30 Rolls Royce Plc Rotor blades
GB0901129D0 (en) 2009-01-26 2009-03-11 Rolls Royce Plc Rotor blade
US8393872B2 (en) * 2009-10-23 2013-03-12 General Electric Company Turbine airfoil
US9004861B2 (en) * 2012-05-10 2015-04-14 United Technologies Corporation Blade tip having a recessed area
DE102013020826A1 (de) * 2013-12-17 2015-06-18 Man Diesel & Turbo Se Radialverdichterstufe
DE102014003123A1 (de) * 2014-03-03 2015-09-03 Mtu Friedrichshafen Gmbh Verdichter
CN105571558B (zh) * 2015-12-30 2018-06-05 中国航空工业集团公司沈阳发动机设计研究所 一种压气机叶片展向波纹度的定量检测方法
CN108506047A (zh) * 2018-03-15 2018-09-07 哈尔滨工业大学 具有三维曲面叶顶的轴流式涡轮叶栅

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GB733918A (en) * 1951-12-21 1955-07-20 Power Jets Res & Dev Ltd Improvements in blades of elastic fluid turbines and dynamic compressors
GB1262182A (en) * 1968-05-12 1972-02-02 Mo Energeticheskij Institut Improvements in or relating to turbine rotor blades
DE19904229A1 (de) * 1999-02-03 2000-08-10 Asea Brown Boveri Gekühlte Turbinenschaufel
US20020182074A1 (en) * 2001-05-31 2002-12-05 Bunker Ronald Scott Film cooled blade tip
US6565324B1 (en) * 1999-03-24 2003-05-20 Abb Turbo Systems Ag Turbine blade with bracket in tip region
EP1510652A2 (fr) * 2003-08-28 2005-03-02 General Electric Company Procédés et appareil pour réduire les vibrations induites des aubes du compresseur

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JPS60184905A (ja) * 1984-03-02 1985-09-20 Mitsubishi Heavy Ind Ltd 軸流回転機械の動翼
JPH10306702A (ja) * 1997-05-08 1998-11-17 Mitsubishi Heavy Ind Ltd ガスタービン翼
US6224336B1 (en) * 1999-06-09 2001-05-01 General Electric Company Triple tip-rib airfoil
JP2002048095A (ja) * 2000-08-03 2002-02-15 Hitachi Ltd 軸流圧縮機の翼
JP2002213206A (ja) * 2001-01-12 2002-07-31 Mitsubishi Heavy Ind Ltd ガスタービンにおける翼構造

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB733918A (en) * 1951-12-21 1955-07-20 Power Jets Res & Dev Ltd Improvements in blades of elastic fluid turbines and dynamic compressors
GB1262182A (en) * 1968-05-12 1972-02-02 Mo Energeticheskij Institut Improvements in or relating to turbine rotor blades
DE19904229A1 (de) * 1999-02-03 2000-08-10 Asea Brown Boveri Gekühlte Turbinenschaufel
US6565324B1 (en) * 1999-03-24 2003-05-20 Abb Turbo Systems Ag Turbine blade with bracket in tip region
US20020182074A1 (en) * 2001-05-31 2002-12-05 Bunker Ronald Scott Film cooled blade tip
EP1510652A2 (fr) * 2003-08-28 2005-03-02 General Electric Company Procédés et appareil pour réduire les vibrations induites des aubes du compresseur

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
OKAPUU U: "AERODYNAMIC DESIGN OF FIRST STAGE TURBINES FOR SMALL AERO ENGINES", LECTURE SERIES - VON KARMANN INSTITUTE FOR FLUID DYNAMICS, RHODE-SAINT-GENESE, BE, 15 June 1987 (1987-06-15), pages 1 - 5, XP002901080, ISSN: 0377-8312 *

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8632311B2 (en) 2006-08-21 2014-01-21 General Electric Company Flared tip turbine blade
DE102009036406A1 (de) * 2009-08-06 2011-02-10 Mtu Aero Engines Gmbh Schaufelblatt
US9011081B2 (en) 2009-08-06 2015-04-21 Mtu Aero Engines Gmbh Blade
US9593584B2 (en) 2012-10-26 2017-03-14 Rolls-Royce Plc Turbine rotor blade of a gas turbine
US10641107B2 (en) 2012-10-26 2020-05-05 Rolls-Royce Plc Turbine blade with tip overhang along suction side
DE102012021400A1 (de) * 2012-10-31 2014-04-30 Rolls-Royce Deutschland Ltd & Co Kg Turbinenrotorschaufel einer Gasturbine
US10267330B2 (en) 2014-06-25 2019-04-23 Siemens Aktiengesellschaft Compressor aerofoil and corresponding compressor rotor assembly
US10107108B2 (en) 2015-04-29 2018-10-23 General Electric Company Rotor blade having a flared tip

Also Published As

Publication number Publication date
EP1591624A1 (fr) 2005-11-02

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