EP1723339B1 - Compresseur d'une turbine a gaz ainsi que turbine a gaz - Google Patents

Compresseur d'une turbine a gaz ainsi que turbine a gaz Download PDF

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Publication number
EP1723339B1
EP1723339B1 EP05715048A EP05715048A EP1723339B1 EP 1723339 B1 EP1723339 B1 EP 1723339B1 EP 05715048 A EP05715048 A EP 05715048A EP 05715048 A EP05715048 A EP 05715048A EP 1723339 B1 EP1723339 B1 EP 1723339B1
Authority
EP
European Patent Office
Prior art keywords
sweep angle
compressor
blade
rotor
blades
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.)
Expired - Fee Related
Application number
EP05715048A
Other languages
German (de)
English (en)
Other versions
EP1723339A1 (fr
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 EP1723339A1 publication Critical patent/EP1723339A1/fr
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Publication of EP1723339B1 publication Critical patent/EP1723339B1/fr
Expired - Fee Related legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D21/00—Pump involving supersonic speed of pumped fluids
    • 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
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades

Definitions

  • the invention relates to a compressor of a gas turbine, in particular an aircraft engine, according to the preamble of claim 1. Furthermore, the invention relates to a gas turbine, in particular an aircraft engine, according to the preamble of claim 4.
  • a compressor is for example from the US-A-2738950 or the EP-A-1 111 188 known.
  • Gas turbines such as aircraft engines, consist of several components, namely a fan or fan, preferably a plurality of compressors, a combustion chamber, and preferably a plurality of turbines. To improve the efficiency and operating range of such gas turbines, it is necessary to optimize all subsystems or components of the gas turbine.
  • the present invention relates to the improvement of the efficiency or working range of compressors, in particular transonic high-pressure compressors.
  • Compressors of gas turbines are usually made of a plurality of axially successively arranged in the flow stages, each stage is formed by associated by a rotor, forming a blade ring blades and a vane ring.
  • the blades associated with the rotor and forming the blade ring, together with the rotor, rotate with respect to the stationary vanes and a housing which is also stationary.
  • To reduce manufacturing costs increasingly compact designs of compressors with the lowest possible number of stages are sought.
  • the overall pressure conditions within the gas turbine or the compressor and thus the step pressure ratios between individual stages increase.
  • the document US 2,738,950 A relates to a turbomachine, in particular a gas turbine whose blading is provided for particularly high gas velocities.
  • the blades have triangular, sharp-edged airfoils.
  • the leading and trailing edges of the airfoils are contoured in a zigzag shape with a forward sweep angle at the radially outer blade tip, followed by a backward sweep angle to the rotor hub, then a forward sweep angle, etc., in the form of a zigzag line.
  • the different fence angles are connected by sharp-edged creases.
  • Such a blade design has never been successful in practice, inter alia because of aerodynamic disadvantages.
  • the present invention is based on the problem to provide a novel compressor of a gas turbine and a novel gas turbine.
  • the efficiency and the working range of the compressor is optimized by the inventive design of the leading edge of the blades.
  • the inventive design of the leading edge of the blades results in an aerodynamically optimal position of a shock wave or shock front of the compressor with respect to the front edge of the impinged blade. It is a recognition of the present invention that the location of the shock wave of the compressor with respect to the leading edge of the blades is important for providing optimum efficiency and operating range of the compressor.
  • the known from the prior art sweeps the leading edges of fan blades only affect the position of a shock or shock wave on a suction side of the fan blades. It is therefore a finding of the present invention that the inventive design of the leading edges of compressor blades an optimized position of the shock front with respect to the leading edge, namely a concern the shock front at the front edge in the radially outer region, can be achieved.
  • the gas turbine according to the invention is defined in claim 4.
  • Fig. 1 shows a section of a compressor 10 according to the invention in the range of two blades 11 and 12 in a view from radially outside.
  • Fig. 2 In addition to the blades 11 and 12 and a rotor hub 13 of the compressor 10 can be seen.
  • the blades 11 and 12 rotate together with the rotor along the direction visualized by the arrow 14.
  • An arrow 15 visualizes the direction of flow or direction of flow of the blades formed by the blades 11 and 12 of the compressor 10.
  • the flow of the blade or the blades 11 and 12 takes place preferably in supersonic region, the outflow of the blades 11 and 12 in the subsonic area he follows.
  • Each of the rotor blades 11 and 12 of the rotor blade grid is essentially delimited by a flow inlet edge or front edge 16, a flow outlet edge or trailing edge 17 and a blade surface 20 forming a suction side 18 and a pressure side 19 between the front edge 16 and the trailing edge 17.
  • the flow of the blades 11 and 12 in the region of the leading edges 16 is preferably in the supersonic range, the outflow thereof in the region of the trailing edges 17 is preferably carried out in the subsonic region.
  • the leading edges 16 of the blades 11 and 12 are designed so that a gas-dynamic compatibility of the blades 11 and 12 is given with a compressor surge.
  • a shock front of the compressor shock in the region of the leading edge 16 of the impinged blade 11 at.
  • a shock front 21 which rests in inventively designed blades on the front edge 16 of the impinged blade 11, in a radially outer region of the leading edge 16.
  • the reference numeral 22 is in Fig. 1 a spaced from the front edge 16 of the impinged blade 11 or detached shock front shown in prior art known compressors, the blades are not formed in accordance with the invention.
  • Such a front-end 16 of the impinged blade 11 detached shock front of the compressor impeller is avoided with the present invention, and thereby the efficiency and operating range of the compressor 10 is optimized.
  • the leading edges 16 of the blades 11, 12 are inclined by a sweeping angle that varies with the height of the blades such that the leading edges 16 at least one forward sweep angle in a radially outward region thereof adjoins the forward sweep angle radially outward Rear sweep angle or zero sweep angle and have a forward sweep angle adjoining the backward sweep angle or zero sweep angle radially outward.
  • This area is in Fig. 2 , which shows a view perpendicular to the suction side 18 of the blade 11, indicated by the reference numeral 23.
  • the suction side 18 of the blade 11 is for reasons of clarity in FIG Fig. 2 hatched, whereas the suction side 18 of the positioned behind the blade 12 is partially hidden by the blade 11 and shown non-hatched.
  • the radially outward region 23 of the leading edges 16 of the blades, in which they have at least the forward sweep angle, the forward sweep angle or zero sweep angle at the forward sweep angle, and the forward sweep angle at the back sweep angle or zero sweep angle, is between 60% and 100%. Preferably, this range is between 70% and 100% of the radial height of the rotor blades 11, 12.
  • the contouring of the leading edges 16 of the blades 11, 12 according to the invention therefore relates to the area of the blade tips of the rotor blades 11, 12 - And, starting from the hub portion 13, the last 40% and 30% of the blades 11, 12th
  • an embodiment of the rotor blades is preferred in which they have two sections with forward sweep angles within the radially outer region 23, wherein a section with a backward sweep angle is positioned between these two sections with a forward sweep angle.
  • the terms forward sweep angle and reverse sweep angle shall be defined such that a blade 11, 12 at a leading edge 16 then has a forward sweep angle at a certain radial height when a point on the leading edge 16 of a blade cut is opposite that radial height the leading edge points is positioned on the hub side adjacent or radially below adjacent or radially within adjacent blade sections upstream.
  • there is a backward sweep angle when a point on the leading edge 16 of the blade section is positioned at a certain radial height with respect to the leading edge points adjacent the hub adjacent radially downstream of radially adjacent blade cuts, respectively.
  • adjacent leading edge points are not fluidly aligned with each other.
  • the flow direction is in Fig. 1 and 3 visualized by an arrow 24.
  • the sweep angle refers to the actual direction of flow of the blade.
  • the leading edge 16 of the blades 11, 12 at a height of about 60% to 80% of the radial height of the blade 11, 12 over a forward sweep angle.
  • this forward sweep angle is at a height of approximately 75% of the radial height of the blade 11, 12.
  • This forward sweep angle is then followed by an area having a backward sweep angle or zero sweep angle, with the leading edge 16 having that backward sweep angle or null sweep angle at a height of about 80% to 90%, and more preferably at a radial height of about 85%.
  • this backward sweep angle or zero sweep angle then in a range at a radial height of approximately 90% to 100%, in turn, a portion of the leading edge 16 adjoins a forward sweep angle.
  • Such a configuration of the rotor blades 11, 12 is particularly preferred gas-dynamically or aerodynamically and ensures a concern of the shock wave of a compression shock to the impinged blade. As a result, the efficiency or working range of the compressor is positively influenced.
  • forward sweep angle and the back sweep angle preferably have values up to 20 °.
  • larger forward sweep angles and backward sweep angles are also possible within the meaning of the invention.
  • the invention relates to the contour of the blade leading edge in the radially outer region 23, which, as already mentioned, between 60% and 100%, preferably between 70% and 100%, of the height of the blade.
  • the region of the blade leading edge 16, which lies between the hub 13 and the radially outer region 23, can be contoured as desired.
  • Fig. 3 schematically different contours of the leading edge 16 in the region between the hub 13 and the invention contoured, radially outer region 23.
  • Fig. 3 in this area between the hub 13 and the radially outer region 23 in dashed lines a backward arrow and with solid lines a forward sweeping the leading edge 16 shown.
  • the contouring of the front edge 16 in the region between the hub 13 and the radially outwardly contoured region 23 according to the invention can be freely selected.
  • the contouring of the trailing edge 17 can be chosen freely.
  • a gas-dynamic and aerodynamically optimized blading of compressor rotors wherein in particular the radially outer blade tips of the blades in the region of the leading edges are designed gas-dynamically compatible with respect to a compressor surge.
  • the head shaft of a compressor shock is applied to the leading edge of the impinged blade.
  • the leading edge of the blade in a radially outer region has at least one hybrid sweep, wherein this hybrid sweep is formed at least by a forward swept portion and a radially outwardly adjoining rearwardly swept portion.
  • the shock front is located on the inventively designed blade in the radially outer, according to the invention contoured region of the front edge of the impinged blade.
  • One Such concerns the shock front of the impinged compressor blade is aerodynamically and gas-dynamically optimal.

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

Abstract

L'invention concerne un compresseur, notamment un compresseur à haute pression, d'une turbine à gaz, notamment d'un groupe motopropulseur. Le compresseur selon l'invention comprend au moins un rotor et plusieurs aubes mobiles (11, 12) associées au rotor ou à chaque rotor et tournant avec leur rotor respectif. Chaque aube mobile (11, 12) est pratiquement délimitée par un bord d'entrée de courant ou bord d'attaque (16), par un bord de sortie de courant ou bord de fuite (17) et par une surface d'ailette (20) s'étendant entre le bord d'attaque (16) et le bord de fuite (17) et formant un extrados (18) et un intrados. Selon l'invention, les bords d'attaque (16) des aubes mobiles (11, 12) sont inclinés d'un angle de flèche variant avec la hauteur de l'aube mobile correspondante (11, 12), de sorte que les bords d'attaque (16) présentent dans leur zone externe radiale (23) au moins un angle de flèche avant, un angle de flèche arrière ou un angle de flèche nulle adjacent de manière radiale extérieure à l'angle de flèche avant et un angle de flèche avant adjacent de manière radiale extérieure à l'angle de flèche arrière ou à l'angle de flèche nulle.

Claims (4)

  1. Compresseur, en particulier compresseur à haute pression, d'une turbine à gaz, en particulier d'un groupe motopropulseur, avec au moins un rotor et plusieurs aubes mobiles (11, 12) associées au rotor ou à chaque rotor et tournant avec leur rotor respectif, chaque aube mobile (11, 12) étant essentiellement délimitée par un bord d'entrée de courant et/ou bord d'attaque (16), par un bord de sortie de courant et/ou bord de fuite (17) et par une surface d'ailette (20) s'étendant entre le bord d'attaque (16) et le bord de fuite (17) et formant un côté d'aspiration (18) et un côté de pression (19), et les bords d'attaque (11) des aubes mobiles (11, 12) étant inclinés d'un angle de flèche variant avec la hauteur de l'aube mobile correspondante (11, 12), de sorte que les bords d'attaque (16) présentent dans leur zone externe radiale (23) un angle de flèche avant, un angle de flèche arrière ou un angle de flèche nulle adjacent de manière radiale extérieure à l'angle de flèche avant et un angle de flèche avant adjacent de manière radiale extérieure à l'angle de flèche arrière ou à l'angle de flèche nulle, caractérisé en ce que les bords d'attaque (16) présentent, avec une hauteur d'environ 60 % à 80 % de la hauteur radiale des aubes mobiles (11, 12), un angle de flèche avant, en ce que les bords d'attaque (16) présentent, avec une hauteur d'environ 80 % à 90 % de la hauteur radiale des aubes mobiles (11, 12), un angle de flèche arrière ou un angle de flèche nulle, et en ce que les bords d'attaque (16) présentent, avec une hauteur d'environ 90 % à 100 % de la hauteur radiale des aubes mobiles (11, 12), un angle de flèche avant.
  2. Compresseur selon la revendication 1, caractérisé en ce que la zone externe radiale (23) des bords d'attaque (16) au niveau de laquelle ceux-ci présentent au moins un angle de flèche avant, un angle de flèche arrière ou un angle de flèche nulle adjacent à l'angle de flèche avant et un angle de flèche avant adjacent à l'angle de flèche arrière ou à l'angle de flèche nulle, se situe à une hauteur comprise entre 65 % et 100 % de la hauteur des aubes mobiles (11, 12).
  3. Compresseur selon la revendication 2, caractérisé en ce que la zone externe radiale (23) des bords d'attaque (16) au niveau de laquelle ceux-ci présentent au moins un angle de flèche avant, un angle de flèche arrière ou un angle de flèche nulle adjacent à l'angle de flèche avant et un angle de flèche avant adjacent à l'angle de flèche arrière ou à l'angle de flèche nulle, se situe à une hauteur comprise entre 70 % et 100 % de la hauteur des aubes mobiles (11, 12).
  4. Turbine à gaz, en particulier groupe motopropulseur, avec au moins un compresseur, en particulier un compresseur à haute pression selon l'une ou plusieurs des revendications 1 à 3.
EP05715048A 2004-03-10 2005-03-03 Compresseur d'une turbine a gaz ainsi que turbine a gaz Expired - Fee Related EP1723339B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102004011607.5A DE102004011607B4 (de) 2004-03-10 2004-03-10 Verdichter einer Gasturbine sowie Gasturbine
PCT/DE2005/000357 WO2005088135A1 (fr) 2004-03-10 2005-03-03 Compresseur d'une turbine a gaz ainsi que turbine a gaz

Publications (2)

Publication Number Publication Date
EP1723339A1 EP1723339A1 (fr) 2006-11-22
EP1723339B1 true EP1723339B1 (fr) 2011-01-26

Family

ID=34962616

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05715048A Expired - Fee Related EP1723339B1 (fr) 2004-03-10 2005-03-03 Compresseur d'une turbine a gaz ainsi que turbine a gaz

Country Status (5)

Country Link
US (1) US7789631B2 (fr)
EP (1) EP1723339B1 (fr)
CA (1) CA2558325A1 (fr)
DE (2) DE102004011607B4 (fr)
WO (1) WO2005088135A1 (fr)

Families Citing this family (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB0620769D0 (en) * 2006-10-19 2006-11-29 Rolls Royce Plc A fan blade
GB0701866D0 (en) 2007-01-31 2007-03-14 Rolls Royce Plc Tone noise reduction in turbomachines
DE102007020476A1 (de) 2007-04-27 2008-11-06 Rolls-Royce Deutschland Ltd & Co Kg Vorderkantenverlauf für Turbomaschinenkomponenten
US20100143138A1 (en) * 2008-12-08 2010-06-10 Russel Hugh Marvin Axial flow wind turbine
KR101810599B1 (ko) 2008-10-30 2017-12-20 파워 제네레이션 테크놀로지스 디베럽먼트 펀드 엘. 피. 환상면 경계층 가스 터빈
US9052116B2 (en) 2008-10-30 2015-06-09 Power Generation Technologies Development Fund, L.P. Toroidal heat exchanger
FR2969230B1 (fr) * 2010-12-15 2014-11-21 Snecma Aube de compresseur a loi d'empilage amelioree
US8684698B2 (en) 2011-03-25 2014-04-01 General Electric Company Compressor airfoil with tip dihedral
US8702398B2 (en) 2011-03-25 2014-04-22 General Electric Company High camber compressor rotor blade
FR2983234B1 (fr) 2011-11-29 2014-01-17 Snecma Aube pour disque aubage monobloc de turbomachine
GB201702383D0 (en) * 2017-02-14 2017-03-29 Rolls Royce Plc Gas turbine engine fan blade with axial lean
US10947851B2 (en) 2018-12-19 2021-03-16 Raytheon Technologies Corporation Local pressure side blade tip lean

Family Cites Families (8)

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Publication number Priority date Publication date Assignee Title
US2738950A (en) * 1945-12-13 1956-03-20 Lockheed Aircraft Corp Turbine machine having high velocity blading
DE1903642A1 (de) * 1969-01-20 1970-08-06 Bbc Sulzer Turbomaschinen Schaufelung fuer Rotoren von Axialverdichtern
US5167489A (en) * 1991-04-15 1992-12-01 General Electric Company Forward swept rotor blade
US5642985A (en) * 1995-11-17 1997-07-01 United Technologies Corporation Swept turbomachinery blade
US6071077A (en) * 1996-04-09 2000-06-06 Rolls-Royce Plc Swept fan blade
US6341942B1 (en) * 1999-12-18 2002-01-29 General Electric Company Rotator member and method
US6328533B1 (en) * 1999-12-21 2001-12-11 General Electric Company Swept barrel airfoil
FR2851798B1 (fr) * 2003-02-27 2005-04-29 Snecma Moteurs Aube en fleche de turboreacteur

Also Published As

Publication number Publication date
CA2558325A1 (fr) 2005-09-22
DE502005010908D1 (fr) 2011-03-10
US7789631B2 (en) 2010-09-07
DE102004011607B4 (de) 2016-11-24
DE102004011607A1 (de) 2005-10-06
EP1723339A1 (fr) 2006-11-22
WO2005088135A1 (fr) 2005-09-22
US20070297904A1 (en) 2007-12-27

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