EP1723339A1 - Compresseur d'une turbine a gaz ainsi que turbine a gaz - Google Patents

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

Info

Publication number
EP1723339A1
EP1723339A1 EP05715048A EP05715048A EP1723339A1 EP 1723339 A1 EP1723339 A1 EP 1723339A1 EP 05715048 A EP05715048 A EP 05715048A EP 05715048 A EP05715048 A EP 05715048A EP 1723339 A1 EP1723339 A1 EP 1723339A1
Authority
EP
European Patent Office
Prior art keywords
sweep angle
height
rotor
compressor
angle
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
EP05715048A
Other languages
German (de)
English (en)
Other versions
EP1723339B1 (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
Application granted granted Critical
Publication of EP1723339B1 publication Critical patent/EP1723339B1/fr
Expired - Fee Related legal-status Critical Current
Anticipated expiration legal-status Critical

Links

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 patent claim 1.
  • the invention further relates to a gas turbine, in particular an aircraft engine, according to the preamble of patent claim 11.
  • Gas turbines such as aircraft engines, consist of several assemblies, namely a fan or fan, preferably several compressors, a combustion chamber, and preferably several 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 operating range of compressors, in particular of transonic high-pressure compressors.
  • Compressors of gas turbines generally consist of several stages arranged axially one behind the other in the flow, each stage being formed by rotor blades assigned to a rotor blade ring and a guide blade ring.
  • the rotor blades assigned to the rotor and forming the rotor blade ring rotate together with the rotor relative to the stationary guide vanes and a housing which is also designed to be stationary.
  • increasingly compact designs of compressors with the lowest possible number of stages are being sought.
  • the total pressure ratios within the gas pressure turbine or the compressor and thus the stage pressure ratios between individual stages increase.
  • stage pressure ratios become increasingly larger and the number of stages becomes increasingly smaller, the circumferential speeds of the rotating components of the compressor are inevitably higher.
  • the speeds increasing with the reduction in the number of stages lead on the one hand to increasing mechanical loads, in particular on the rotor blades rotating with the rotor, and on the other hand to a so-called supersonic flow against the rotor blades and to transonic flow conditions within the blade grids.
  • the present invention is based on the problem of creating a new type of compressor for a gas turbine and a new type of gas turbine.
  • the leading edges of the moving blades are inclined by an arrowing angle that changes with the height of the respective rotating blades such that the leading edges in a radially outer region thereof have at least one forward arrowing angle, a backward arrowing angle or zero arrowing angle that adjoins the forward arrowing angle radially on the outside, and a backward arrowing angle or have the forward sweep angle adjoining the zero sweep angle radially on the outside.
  • the efficiency and the working range of the compressor are optimized by the inventive design of the leading edge of the rotor blades.
  • the inventive design of the front edge of the rotor 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 flowed rotor blade. It is a finding of the present invention that the position of the shock front or shock wave of the compressor with respect to the front edge of the rotor blades is important for providing an optimal efficiency and working range of the compressor.
  • the sweeps of the front edges of fan blades known from the prior art only influence the position of a shock front or shock wave on a suction side of the fan blades.
  • the radially outer region of the leading edges in which they have at least one forward sweep angle, one backward sweep angle or zero sweep angle following the forward sweep angle and one forward sweep angle following the reverse sweep angle or zero sweep angle, lie between 60% and 100%, preferably between 70% and 100%, the height of the blades.
  • the leading edges of the rotor blades in this radially outer region in the direction from radially inward to radially outward have a forward sweep angle, a backward sweep angle following the forward sweep angle and a forward sweep angle following the reverse sweep angle.
  • the radially outer region between 60% and 100%, preferably between 70% and 100%, of the height of the rotor blade two sections which are arrowed forward then include a section which is arrowed backwards.
  • the gas turbine according to the invention is defined in claim 11.
  • FIG. 1 shows a schematic section of a compressor according to the invention in a view from the ' radially outside of two rotor blade profiles of the compressor according to the invention, the two rotor blade profiles being shown in cross section along the section line II at approximately 80% of the blade height according to FIG. 3 ;
  • 2 shows a schematic section of a compressor according to the invention in a view perpendicular to the suction side of a rotor blade of the compressor;
  • FIG. 3 shows a schematic section of a compressor according to the invention in a meridian plane view of the compressor together with the position of a compression joint near the front edge of the rotor blade.
  • the present invention is described below with reference to FIG.
  • FIG. 1 shows a detail from a compressor 10 according to the invention in the area of two rotor blades 11 and 12 in a view from the radial outside.
  • a rotor hub 13 of the compressor 10 can also be seen in FIG. 2.
  • the blades 11 and 12 rotate together with the rotor along the direction visualized by the arrow 14.
  • An arrow 15 visualizes the flow direction or flow direction of the rotor vane grille of the compressor 10 formed by the rotor blades 11 and 12.
  • the flow against the rotor vane grille or the rotor blades 11 and 12 preferably takes place in the supersonic region, the outflow from the rotor blades 11 and 12 in the subsonic region he follows.
  • Each of the rotor blades 11 and 12 of the rotor blade grille is essentially delimited by a flow inlet edge or front edge 16, a flow outlet edge or rear edge 17 and a blade surface 20 which extends between the front edge 16 and the rear edge 17 and forms a suction side 18 and a pressure side 19.
  • the flow against the blades 11 and 12 in the area of the front edges 16 preferably takes place in the supersonic area
  • the outflow thereof in the area of the rear edges 17 preferably takes place in the subsonic area.
  • the front edges 16 of the blades 11 and 12 are designed in such a way that the blades 11 and 12 are gas-dynamically compatible with a compressor shock.
  • the reference number 22 in FIG. 1 shows a shock front which is spaced apart or detached from the front edge 16 of the flow-moving blades 11 and which occurs in compressors known from the prior art, the blades of which are not designed in accordance with the invention.
  • a shock front of the compressor shock, detached in this way from the front edge 16 of the rotor blade 11 against which flow is flowing, is avoided with the present invention, and the efficiency and operating range of the compressor 10 are thereby optimized.
  • the front edges 16 of the blades 11, 12 are inclined by an arrow angle that changes with the height of the blades, such that the front edges 16 in a radially outer region thereof have at least one forward arrow angle, one that adjoins the forward arrow angle radially on the outside Reverse sweep angle or zero sweep angle and have a forward sweep angle radially outward following the reverse sweep angle or zero sweep angle.
  • This area is identified in FIG. 2, which shows a view perpendicular to the suction side 18 of the blade 11, with the reference numerals 23.
  • the suction side 18 of the rotor blade 11 is shown hatched for reasons of a clearer illustration in FIG. 2, whereas the suction side 18 of the rotor blade 12 positioned behind it is partially covered by the rotor blade 11 and is not hatched.
  • the radially outer region 23 of the leading edges 16 of the moving blades, in which they have at least the forward sweep angle, the backward sweep angle or zero sweep angle radially adjacent to the forward sweep angle and the forward sweep angle radially outward following the reverse sweep angle or zero sweep angle, is between 60% and 100% the radial height of the rotor blades 11, 12. This range is preferably between 70% and 100% of the radial height of the rotor blades 11, 12.
  • the contouring of the front edges 16 of the rotor blades 11, 12 according to the invention therefore relates to the region of the blade tips of the rotor blades 11, 12 - namely, starting from the hub area 13, the last 40% or 30% of the blades 11, 12.
  • an embodiment of the rotor blades is preferred in which they have two sections with forward sweep angles within the radially outer region 23, a section with a backward sweep angle being positioned between these two sections with a forward sweep angle.
  • forward sweep angle and backward sweep angle are to be defined such that a blade 11, 12 on a front edge 16 has a forward sweep angle at a certain radial height if a point on the front edge 16 of a blade cut opposite this radial height the leading edge points is positioned upstream on the hub side adjacent or radially below adjacent or radially within adjacent blade sections.
  • there is a backward sweep angle when a point on the leading edge 16 of the blade section is positioned downstream at a certain radial height relative to the leading edge points on the hub side adjacent or radially below adjacent or radially within adjacent blade sections.
  • neighboring leading edge points are not fluidically offset from one another.
  • the flow direction is visualized by an arrow 24 in FIGS. 1 and 3.
  • the sweep angle refers to the actual flow direction of the rotor blade.
  • the front edge 16 of the blades 11, 12 has a forward sweep angle at a height of approximately 60% to 80% of the radial height of the blades 11, 12.
  • This forward sweep angle is approximately 75% of the radial height of the rotor blades 11, 12 is particularly preferred.
  • a region with a backward sweep angle or zero sweep angle then adjoins this forward sweep angle, the front edge 16 having this backward sweep angle or zero sweep angle at a height of approximately 80% to 90%, in particular at a radial height of approximately 85%.
  • This backward-sweeping angle or zero-sweeping angle is then preferably followed in turn by a region of the front edge 16 with a forward sweeping angle in a range at a radial height of approximately 90% to 100%.
  • Such a configuration of the rotor blades 11, 12 is particularly preferred in terms of gas dynamics or aerodynamics and ensures that the shock wave of a compression shock is in contact with the flow against the rotor blade. This has a positive influence on the efficiency or working range of the compressor.
  • the forward sweep angle and the reverse 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 contouring of the blade leading edge in the radially outer region 23, which, as already mentioned, between 50% and 100%, in particular between 60% and 100%, preferably between 70% and 100% , the height of the Laufetzaufei.
  • the area of the blade leading edge 16, which lies between the hub 13 and the radially outer area 23, can be contoured as desired.
  • 3 schematically shows different contours of the front edge 16 in the area between the hub 13 and the radially outer area 23, which is contoured according to the invention.
  • FIG. 3 in this area between the hub 13 and the radially outer area 23 is shown in broken lines Lines show a backward arrow and a solid arrow shows a forward arrow of the front edge 16.
  • the contour of the rear edge 17 can also be freely selected.
  • gas-dynamically and aerodynamically optimized blading of compressor rotors is accordingly provided, in particular the radially outer blade tips of the rotor blades in the region of the leading edges being designed to be gas-dynamically compatible with regard to a compressor shock.
  • the top shaft of a compressor shock rests against the front edge of the rotor blade against which the air flows.
  • the front edge of the rotor blade has at least one hybrid sweep in a radially outer region, this hybrid sweep being formed at least by a forward-swept section and a backward-sweeping section adjoining it radially on the outside.
  • At least the following advantages result: a better efficiency of the compressor is achieved; the compressor has an extended operating range with good efficiency and thus a wider working range; the surge margin of the compressors is optimized; the vibration behavior is improved by the changing radial distribution of the chord length; there is an improved brushing behavior of the moving blades.
  • the butt front bears against the rotor blade designed according to the invention in the radially outer region of the front edge of the flowed rotor blade which is contoured according to the invention. On Such an impact of the impact front on the flow against the compressor blade is optimal in terms of aerodynamics and gas dynamics.

Landscapes

  • 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.
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 true EP1723339A1 (fr) 2006-11-22
EP1723339B1 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
US9052116B2 (en) 2008-10-30 2015-06-09 Power Generation Technologies Development Fund, L.P. Toroidal heat exchanger
CN102203388B (zh) 2008-10-30 2015-11-25 电力技术发展基金公司 环形边界层气体涡轮机
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)

* Cited by examiner, † Cited by third party
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

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO2005088135A1 *

Also Published As

Publication number Publication date
US7789631B2 (en) 2010-09-07
DE502005010908D1 (fr) 2011-03-10
DE102004011607A1 (de) 2005-10-06
WO2005088135A1 (fr) 2005-09-22
DE102004011607B4 (de) 2016-11-24
CA2558325A1 (fr) 2005-09-22
EP1723339B1 (fr) 2011-01-26
US20070297904A1 (en) 2007-12-27

Similar Documents

Publication Publication Date Title
EP1723339A1 (fr) Compresseur d'une turbine a gaz ainsi que turbine a gaz
EP1774179B1 (fr) Aube de compresseur
EP2108784B1 (fr) Turbomachine dotée d'un composant d'injecteur de fluide
EP2242931B1 (fr) Structure d'écoulement pour turbocompresseur
EP2194232B1 (fr) Turbomachine dotée d'une barrière à couche frontière sur la paroi latérale
DE60031941T2 (de) Geneigtes Schaufelblatt mit tonnenförmiger Anströmkante
EP2226509B1 (fr) Turbo compresseur ou pompe avec injection de fluide pour influencer la couche limite
EP1759090A1 (fr) Aube comportant une zone de transition
DE60312466T2 (de) Gasturbinenleitschaufel sowie Gasturbine umfassend eine Gasturbinenleitschaufel
CH710476A2 (de) Verdichter mit einer Axialverdichterendwandeinrichtung zur Steuerung der Leckageströmung in dieser.
DE602004005906T2 (de) Gekrümmte Schaufel für Strahltriebwerk
EP2249044A2 (fr) Compresseur ou pompe avec extraction de fluide
DE102007024840A1 (de) Strömungsarbeitsmaschinenschaufel mit Multi-Profil-Gestaltung
DE102010014556B4 (de) Leitschaufel eines Verdichters
EP2597257B1 (fr) Aubage
EP3715586B1 (fr) Pale d'aube de rotor d'une turbomachine
DE102016124147A1 (de) Innenkühlkonfigurationen in Turbinenrotorschaufeln
DE102019135335A1 (de) Hybrid-laufschaufeln für turbinentriebwerke
DE102016122696A1 (de) Eintrittsleitrad für eine Turbomaschine
DE102019200885A1 (de) Leitgitter für eine Strömungsmaschine
EP2631429A1 (fr) Aubage
DE102022103319A1 (de) Leitschaufel für eine Strömungsmaschine
DE202023106979U1 (de) Leitschaufelcluster für eine Strömungsmaschine
DE102020117944A1 (de) Impellereinheit, Pumpe, Raketentriebwerk und Verfahren zur Herstellung einer Impellereinheit
EP1956190B1 (fr) Profil pour une aube de turbine et installation de turbine

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20060825

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): DE FR GB IT

DAX Request for extension of the european patent (deleted)
RBV Designated contracting states (corrected)

Designated state(s): DE FR GB IT

17Q First examination report despatched

Effective date: 20100511

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB IT

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

REF Corresponds to:

Ref document number: 502005010908

Country of ref document: DE

Date of ref document: 20110310

Kind code of ref document: P

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 502005010908

Country of ref document: DE

Effective date: 20110310

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed

Effective date: 20111027

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 502005010908

Country of ref document: DE

Effective date: 20111027

REG Reference to a national code

Ref country code: GB

Ref legal event code: 732E

Free format text: REGISTERED BETWEEN 20131114 AND 20131120

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 12

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 13

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 14

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20200325

Year of fee payment: 16

Ref country code: DE

Payment date: 20200324

Year of fee payment: 16

Ref country code: IT

Payment date: 20200325

Year of fee payment: 16

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20200325

Year of fee payment: 16

REG Reference to a national code

Ref country code: DE

Ref legal event code: R119

Ref document number: 502005010908

Country of ref document: DE

GBPC Gb: european patent ceased through non-payment of renewal fee

Effective date: 20210303

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: DE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20211001

Ref country code: FR

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20210331

Ref country code: GB

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20210303

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: IT

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20210303