EP3762587A1 - Pale d'aube pour une aube de turbine - Google Patents

Pale d'aube pour une aube de turbine

Info

Publication number
EP3762587A1
EP3762587A1 EP19723730.8A EP19723730A EP3762587A1 EP 3762587 A1 EP3762587 A1 EP 3762587A1 EP 19723730 A EP19723730 A EP 19723730A EP 3762587 A1 EP3762587 A1 EP 3762587A1
Authority
EP
European Patent Office
Prior art keywords
blade
rows
airfoil
cooling holes
height
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
EP19723730.8A
Other languages
German (de)
English (en)
Other versions
EP3762587B1 (fr
Inventor
Fathi Ahmad
Daniela Koch
Marco Schüler
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.)
Siemens Energy Global GmbH and Co KG
Original Assignee
Siemens AG
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 AG filed Critical Siemens AG
Publication of EP3762587A1 publication Critical patent/EP3762587A1/fr
Application granted granted Critical
Publication of EP3762587B1 publication Critical patent/EP3762587B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the invention relates to an airfoil for a turbine blade, comprising a vorström of a hot gas vordarkante from which a suction side wall and a
  • Such a turbine blade is known, for example, from EP 2 154 333 A2.
  • the arranged in the leading edge cooling holes are used during operation of a crudestat ended gas turbine to produce a cooling protective film over the leading edge to counter the incoming hot gas flow ent.
  • the cooling holes are therefore also referred to as film cooling holes, which are also known in English because of their th th arrangement also known as "Shower Head Film Cooling Holes.”
  • film cooling holes which are also known in English because of their th th arrangement also known as "Shower Head Film Cooling Holes.”
  • the stagnation point since there is no cross flow in the idealized Sin ne, for this reason, in the prior art on both sides of the leading edge or the previously determined stagnation film cooling holes arranged in order to prevent the hot gas flow impinging there does not come in too close con tact with the component wall.
  • the disadvantage is that the stagnation point of a show felprofils or the stagnation of a blade can be dependent on different factors, so that there is a
  • US 2016/0010463 A1 teaches, with a displacement of the stagnation line, to arrange on the pressure side of moving blades an additional half-row of film cooling holes on the radially outer half of the blade.
  • the additional film cooling holes increase the consumption of cooling air, which has a negative effect on the efficiency of a turbine equipped therewith.
  • adapted cooling can also be achieved by not selecting the position but merely the inclination of some leading edge film cooling holes in a previously determined displacement of the stagnation dot line so that they do not converge in the opposite direction to the expected local hot gas flow Direction the cooling air ausbla sen, but in the same direction.
  • the present invention seeks to provide an airfoil for a turbine blade, which is best possible decor with different operating conditions of a gas turbine, especially when using a reasonable amount of coolant sufficient cooling with the highest possible life of the airfoil to achieve.
  • the invention is based on the finding that the fact neuter hot gas flow direction of the for the interpretation of Blade sheet used flow direction may differ on the one hand due to different operations of the gas turbine. The deviations can occur due to a change in load to nominal load.
  • the stagnation point of a blade profile in the region of the leading edge can oscillate due to flow effects which are caused by a vane arranged upstream of the blade. The oscillation of the stagnation point of a blade profile leads to locally increased surface temperature of the blade, which can be effectively counteracted by the invention.
  • cooling holes are displaced towards the pressure side or suction side, based on the oscillating stagnation point of the relevant blade profile.
  • an area is determined for each blade profile in which the stagnation point can occur.
  • Each of these areas is defined by two endpoints, from which a mean congestion point can then be determined.
  • the two cooling holes are positioned so that the best possible cooling is achieved. This optimizes the cooling effect locally.
  • By using only two rows of cooling instead of usually three or more complete rows of cooling can also reduce the amount of coolant required for cooling. The reduced consumption of coolant contributes during the operation of the gas turbine to its effect degree increase.
  • the at least two rows of cooling holes along the Ge entire extension of the leading edge between 0% and 100% show blade height on a wavy line with several troughs and wave crests are arranged. Consequently, the cooling holes of the at least two rows are repeatedly displaced slightly locally to the pressure side, compared with cooling holes on another vane blade height.
  • the at least two rows of cooling holes are arranged only partially along the leading edge on a wavy line, such that the at least two rows of cooling holes in a ers th area, which is located between 0% and about 40% airfoil height are arranged substantially parallel on both sides of the leading edge and arranged in a immediacy bar adjacent second region, which extends between about 40% and about 75% vane height and higher, are arranged on the pressure side and wherein the at least two rows of cooling holes in a to the second Be rich immediately adjacent third area, which ends at 100% blade height, with rising Schaufelblatthö hey are arranged back to the front edge back.
  • This refinement is based on the finding that the displacement of the stagnation point of a blade profile in the radially inner region of the blade blade is rather narrow-banded, whereas from a blade blade height of about 40% the displacement increases and, moreover, is more on the pressure side. Accordingly, the cooling holes of the at least two rows are shifted in the range of 40% to 100% to the pressure side, wherein preferably at about 75% blade height, the maxima le pressure-side displacement is arranged. Based on a chord length of the airfoil, the value of the maximum pressure displacement is not more than 5% of the blade pitch, but preferably at least 2%.
  • the first distance between the at least two rows of cooling holes along the leading edge vari, so that the first distance for some blade heights is different. With this measure, the local cooling capacity of the turbine blade in the region of the leading edge to the individual temperature load locally adapted to who.
  • a blade profile can be determined for each blade height by a cross-sectional view wel Ches known to have the shape of a curved drop on.
  • Each blade profile therefore has a nose radius in the region of the front edge, wherein the blade profiles at the height of cooling holes have a first distance between the at least two rows whose size is in the range between 0.4 times and 0.7 times the associated nose radius is.
  • the effectiveness of the cooling depends on the distance between the cooling holes under different rows and the curvature of the leading edge, the so-called nose radius and the length of the camber line, the number of blades and the turning of the blade profile.
  • the ers te distance at half blade height is the smallest and increases towards the two ends. The increase is particularly moderate.
  • each measure cooling hole has a coolant flow SETTING sirloin throttle cross-section, wherein the throttle cross-sections of some cooling holes are different sizes.
  • Particularly preferred are the throttle cross-sections of the cooling holes in Be rich half the blade height greater than the throttle cross-section of the cooling holes in half of Schaufelblatthö he far farther.
  • the embodiment in which the at least two rows of cooling holes are arranged on both sides of a medium ren stagnation point line of the incoming hot gas flow.
  • the hot gas flow is divided into a part to the pressure side and flowing to the suction side dividing part on both sides deflected so that due to the two-sided arrangement of the cooling holes, the underlying component wall is particularly efficiently protected from the high Tempe temperatures of the hot gas.
  • the cooling holes of each of the at least two rows near the fußseiti gene end and near the tip end of the blade are further arranged on the suction side as the cooling holes of the corre sponding row at half blade height.
  • the wavy line then extends between these points without any change. tion of the sign of its curvature, so that it is only ge slightly curved. In-depth investigations have shown that this variant represents a more favorable cooling configuration, in particular for guide vanes, since with these vanes the stagnation point displacement occurs at the ends of the blade blade rather than in its center and also towards the suction side.
  • the maximum displacement of the respective cooling holes near the ends of the airfoil is then only ei few millimeters, in particular 2 mm, towards the suction side, ver compared with the position of the cooling holes of the same row at half the blade height, ie at 50% of the blade height.
  • the at least two rows are provided with a further, but shortened row of essentially uniformly spaced cooling holes, the length of the another row between 50% and 60% of the blade height and the further row of cooling holes is arranged substantially centrally between the two ends of the airfoil.
  • the wei tere series is in the meaning of this application as long as in Wesentli chen arranged centrally, as long as it is divided by the half show felblatt Love in two parts whose shorter part is not shorter than 1/3 of the length of the other row.
  • the length of the further row of cooling holes is detected in the same direction as the blade height.
  • Preferred dimensions of the blade is part of a turbine blade, in particular a turbine vane of a sta tionary gas turbine.
  • FIG. 1 is a perspective view of a turbine blade with an inventive blade according to a first embodiment
  • FIG. 2 shows a perspective view of a turbine blade with an inventive blade according to a second embodiment
  • Figure 4 is a perspective view of a Turbinenleit blade with an airfoil according to the invention according to a third embodiment.
  • a turbine blade 10 is shown in perspective view.
  • the turbine blade 10 comprises in succession a substantially fir-tree-shaped blade root 12, to which a hot gas platform 14 adjoins as the end wall.
  • an inventive blade 16 is arranged according to a first embodiment.
  • the airfoil 16 includes known a leading edge 18 and a trailing edge 20, between which a suction side wall 17 and a Drucksei tenwand 19 extends. In a transverse direction, the airfoil 16 extends from a root end 21 at 0% airfoil height to a tip end 23 at 100% airfoil height.
  • two rows Ri, R2 of cooling holes 22 are arranged.
  • the two Rei hen Ri, R2 run along a wavy line with several Wave troughs and wave crests and are arranged on both sides of a mean stagnation point line 24 simultaneously.
  • FIG. A second embodiment of the invention is shown in FIG.
  • a region is rectilinear, followed by a bulbous portion.
  • the two rows Ri, R2 of cooling holes 22 in the ers th, radially inner region are arranged so that they are arranged parallel to the front edge 18 on both sides thereof.
  • This first area Bi extends between 0% and about 40% vane height.
  • a second region B2 is provided. This ends at a blade height of about 75%.
  • the cooling holes 22 of both rows Ri, R2 shift with increasing height in the direction of the pressure side until they have reached the maximum displacement of the leading edge 18 at about 75% blade height.
  • the cooling holes 22 of the two rows Ri, R2 shift back in the direction of the front edge 18.
  • cooling holes 22 are shown only schematically as circles, the throttle cross-sections schematically represented by different sized circles have been.
  • the cooling holes 22 may be film cooling holes having a diffuser-like opening. Their diffuser can even be profiled from designed. Also, on the surface of the blade 16 transversely to be detected distance A between thedelö Chern 22 may be at different blade heights under different sizes.
  • Figure 3 also shows as a blade profile 28 the cross section through the airfoil 16 of the first embodiment shown in FIG 1.
  • the blade profile center line is provided with the reference numeral 30 Be.
  • the vorderst arranged point of the blade profile center line 30 defines the leading edge 18.
  • the stagnation point 25 off the front edge 18 towards the pressure side 19 or towards the suction side 17 may be slightly shifted.
  • the (average) Stagnationspunk te 25 each blade profile section, which can be determined on any show blade heights together form the congestion point line 24.
  • the nose radius is denoted by R.
  • FIG. It shows in perspective a turbine blade configured as a guide blade, wherein the blade root 12 comprises two hook-shaped rails for fastening the blade to a blade carrier (not shown further).
  • a platform 14 is provided for limiting the flow path both at the foot-side end 21 and at the tip-side end 23 of the airfoil. In between, the blade 16 extends along its blade height.
  • the at least two Rei hen Ri, R2 of cooling holes 18 are arranged analogously: starting with the cooling holes on the half blade height are within each row Ri, R2 arranged with decreasing distance to the platforms 14 down cooling holes further suction side.
  • the stagnation dot line 24 is slightly curved without a change in the sign of its curvature.
  • a further, but shortened series of substantially uniformly spaced cooling holes 18 is provided on the pressure side next to the two rows Ri, R2.
  • This further Rei hey R3 is according to this embodiment, centrally between the two platforms 14 and the two ends 21, 23 is arranged and extends only over a length of 55% of the blade height. It is thus shorter than the two rows Ri, R2. If necessary, locally further, unique cooling holes can be provided near the leading edge.
  • the invention relates to an airfoil 16 for ei ne turbine blade 10, comprising a vortex of a hot gas S front edge 18, from which a suction side wall 17 and a pressure side wall 19 extend to a trailing edge 20 of the blade 16, wherein the airfoil 16 in a Transverse direction extends from a foot-side end 21 with a blade height of 0% to a tip-side end 23 with a blade height of 100%, with two arranged along the leading edge rows Ri, R2 of cooling holes 22 to each other a perpendicular to the Vorderkan te 18 to have sensing first distance A.

Abstract

L'invention concerne une aube (16) pour une aube de turbine, comprenant un bord d'attaque (18) à afflux d'un gaz chaud (S) et à partir duquel une paroi latérale d'aspiration (17) et une paroi latérale de pression (19) s'étendent vers un bord arrière (20) de l'aube (16), l'aube (16) s'étendant dans une direction transversale à partir d'une extrémité de pied (21) ayant une hauteur d'aube de 0% jusqu'à une extrémité de pointe (23) ayant une hauteur d'aube de 100%, ayant au moins deux rangées (R1, R2) de trous de refroidissement (22) disposés le long du bord avant et ayant une première distance (A) à prendre en compte perpendiculairement au bord avant. L'invention vise à fournir une aube de turbine qui, avec une dépense de refroidissement réduite, est capable de refroidir de manière fiable le bord d'attaque (18) pour différentes conditions de fonctionnement. À cet effet, les au moins deux rangées (R1, R2) de trous de refroidissement (22) sont disposées au moins partiellement le long du bord d'attaque (18) sur une ligne ondulée.
EP19723730.8A 2018-05-04 2019-05-03 Pale d'aube pour une aube de turbine Active EP3762587B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP18170731.6A EP3564483A1 (fr) 2018-05-04 2018-05-04 Pale d'aube pour une aube de turbine
PCT/EP2019/061354 WO2019211427A1 (fr) 2018-05-04 2019-05-03 Pale d'aube pour une aube de turbine

Publications (2)

Publication Number Publication Date
EP3762587A1 true EP3762587A1 (fr) 2021-01-13
EP3762587B1 EP3762587B1 (fr) 2022-04-13

Family

ID=62116325

Family Applications (2)

Application Number Title Priority Date Filing Date
EP18170731.6A Withdrawn EP3564483A1 (fr) 2018-05-04 2018-05-04 Pale d'aube pour une aube de turbine
EP19723730.8A Active EP3762587B1 (fr) 2018-05-04 2019-05-03 Pale d'aube pour une aube de turbine

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP18170731.6A Withdrawn EP3564483A1 (fr) 2018-05-04 2018-05-04 Pale d'aube pour une aube de turbine

Country Status (6)

Country Link
US (1) US11326458B2 (fr)
EP (2) EP3564483A1 (fr)
JP (1) JP7124122B2 (fr)
KR (1) KR102505046B1 (fr)
CN (1) CN112074652B (fr)
WO (1) WO2019211427A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3564483A1 (fr) * 2018-05-04 2019-11-06 Siemens Aktiengesellschaft Pale d'aube pour une aube de turbine
JP7224928B2 (ja) * 2019-01-17 2023-02-20 三菱重工業株式会社 タービン動翼及びガスタービン
KR102507408B1 (ko) 2022-11-11 2023-03-08 터보파워텍(주) 3d프린팅에 의한 가스터빈 블레이드의 에어포일 수리 공정

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6869268B2 (en) * 2002-09-05 2005-03-22 Siemens Westinghouse Power Corporation Combustion turbine with airfoil having enhanced leading edge diffusion holes and related methods
US7217094B2 (en) * 2004-10-18 2007-05-15 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
EP1898051B8 (fr) * 2006-08-25 2017-08-02 Ansaldo Energia IP UK Limited Aube de turbine à gaz avec refroidissement du bord d'attaque
US8105030B2 (en) 2008-08-14 2012-01-31 United Technologies Corporation Cooled airfoils and gas turbine engine systems involving such airfoils
WO2014025571A1 (fr) * 2012-08-06 2014-02-13 General Electric Company Composant de turbine rotatif à alignement de trou préférentiel
EP2964932B1 (fr) * 2013-03-04 2020-11-04 United Technologies Corporation Aube et moteur à turbine à gaz associé
US10329923B2 (en) * 2014-03-10 2019-06-25 United Technologies Corporation Gas turbine engine airfoil leading edge cooling
US9976423B2 (en) * 2014-12-23 2018-05-22 United Technologies Corporation Airfoil showerhead pattern apparatus and system
US10240462B2 (en) * 2016-01-29 2019-03-26 General Electric Company End wall contour for an axial flow turbine stage
US11286787B2 (en) * 2016-09-15 2022-03-29 Raytheon Technologies Corporation Gas turbine engine airfoil with showerhead cooling holes near leading edge
EP3564483A1 (fr) * 2018-05-04 2019-11-06 Siemens Aktiengesellschaft Pale d'aube pour une aube de turbine

Also Published As

Publication number Publication date
CN112074652A (zh) 2020-12-11
KR20210002709A (ko) 2021-01-08
WO2019211427A1 (fr) 2019-11-07
JP7124122B2 (ja) 2022-08-23
EP3762587B1 (fr) 2022-04-13
CN112074652B (zh) 2023-05-02
KR102505046B1 (ko) 2023-03-06
JP2021522444A (ja) 2021-08-30
EP3564483A1 (fr) 2019-11-06
US11326458B2 (en) 2022-05-10
US20210156263A1 (en) 2021-05-27

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