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EP3564483A1 - Pale d'aube pour une aube de turbine - Google Patents

Pale d'aube pour une aube de turbine Download PDF

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Publication number
EP3564483A1
EP3564483A1 EP18170731.6A EP18170731A EP3564483A1 EP 3564483 A1 EP3564483 A1 EP 3564483A1 EP 18170731 A EP18170731 A EP 18170731A EP 3564483 A1 EP3564483 A1 EP 3564483A1
Authority
EP
European Patent Office
Prior art keywords
blade
cooling holes
rows
airfoil
leading edge
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Withdrawn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
EP18170731.6A
Other languages
German (de)
English (en)
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 AG
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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, Siemens Corp filed Critical Siemens AG
Priority to EP18170731.6A priority Critical patent/EP3564483A1/fr
Priority to US17/048,582 priority patent/US11326458B2/en
Priority to PCT/EP2019/061354 priority patent/WO2019211427A1/fr
Priority to KR1020207034682A priority patent/KR102505046B1/ko
Priority to CN201980030091.4A priority patent/CN112074652B/zh
Priority to EP19723730.8A priority patent/EP3762587B1/fr
Priority to JP2020561773A priority patent/JP7124122B2/ja
Publication of EP3564483A1 publication Critical patent/EP3564483A1/fr
Withdrawn legal-status Critical Current

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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
    • F01D5/18—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 front edge, which can be flowed on by a hot gas, from which a suction side wall and a pressure side wall extend to a trailing edge of the airfoil, the airfoil extending in a transverse direction from a base end with a blade air height of 0 % to a tip-side end having a blade height of 100%, with two rows of cooling holes arranged along the leading edge, which have a first distance to be detected perpendicular to the leading edge relative to one another.
  • Such a turbine blade is for example from the EP 2 154 333 A2 known.
  • the cooling holes disposed in the leading edge serve to provide a cooling protective film over the leading edge during operation of a gas turbine equipped therewith to counteract the incoming hot gas flow.
  • the cooling holes are therefore also known as film cooling holes, which are also known in English because of their tight arrangement also known as "Shower Head Film Cooling Holes".
  • the airfoil divides the inflowing hot gas flow at the leading edge into two partial flows, of which one partial flow flows along the suction side of the airfoil and the other part along the pressure side.
  • the location of the flow distribution on the blade profile is called the stagnation point, since in the idealized sense there is no cross flow.
  • film cooling holes are arranged in the prior art on both sides of the leading edge or the previously determined stagnation line in order to prevent the hot gas flow impinging there from coming into close contact with the component wall.
  • the present invention seeks to provide an airfoil for a turbine blade, which is best suited for different operating conditions of a gas turbine, in particular to achieve sufficient cooling with the highest possible life of the airfoil when using a reasonable amount of coolant ,
  • the invention is based on the finding that the actual hot gas flow direction can deviate from the flow direction used for the design of the airfoil on the one hand due to different operating modes of the gas turbine. The deviations may occur due to a change in the rated load load.
  • the stagnation point of a blade profile can oscillate in the region of the leading edge due to flow effects 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.
  • a range 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 stagnation point can 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 rows of cooling can also reduce the amount of coolant required for cooling. The reduced consumption of coolant contributes during operation of the gas turbine to increase its efficiency.
  • the two rows of cooling holes along the entire extension of the leading edge between 0% and 100% blade height are arranged on a wavy line with multiple troughs and wave crests.
  • the cooling holes of the two rows are repeatedly shifted locally to the pressure side slightly, compared to cooling holes on a different blade height.
  • the two rows of cooling holes are only partially along the leading edge on a wavy line, such that the two rows of cooling holes in a first region, which is located between 0% and about 40% blade height, substantially parallel on both sides the leading edge are arranged and arranged in a second region immediately adjacent thereto, which extends between about 40% and about 75% blade height and higher, arranged on the pressure side and wherein the two rows of cooling holes in a third region immediately adjacent to the second region, which ends at 100% blade height, are arranged with increasing blade blade height shifted back to the front edge.
  • 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 two rows are shifted in the range of 40% to 100% to the pressure side, wherein preferably at about 75% blade height, the maximum pressure-side displacement is arranged. Based on a chord length of the airfoil, the value of the pressure-side maximum displacement is not more than 5% of the blade chord length, but preferably at least 2% minimum.
  • the first distance between the two rows of cooling holes varies along the front edge, so that the first distance is different for some blade blade heights.
  • the local cooling capacity of the turbine blade in the region of the leading edge can be adapted locally to the individual temperature load.
  • a blade profile can be determined for each blade height by a cross-sectional view, which is known to have the shape of a curved drop.
  • Each blade profile therefore has a nose radius in the region of the front edge, wherein the blade profiles at the level of cooling holes have a first distance between the two rows whose size is in the range between 0.4 times and 0.7 times the associated nose radius .
  • the effectiveness of the cooling depends on the distance of the cooling holes of different rows and the curvature of the leading edge, the so-called nose radius and the length of the camberline, the number of blades and the turning of the blade profile. It was then found that a particularly efficient cooling of the leading edge region can be achieved if the first distance between the lying on the same blade height cooling holes of different rows in the claimed interval.
  • the first distance at half blade height is the smallest and increases toward the two ends.
  • the increase is particularly moderate.
  • each cooling hole preferably has a throttle cross section which adjusts the coolant flow, the throttle cross sections of some cooling holes being of different sizes.
  • the throttle cross-sections of the cooling holes in the region of half the blade blade height are greater than the throttle cross-section of the cooling holes in the farther from half the blade blade height range.
  • This embodiment is based on the finding that at half the blade blade height and the areas immediately adjacent to it, a somewhat higher cooling requirement prevails than in those areas of the leading edge which are farther from the half blade height.
  • the two rows of cooling holes are arranged on both sides of a mean stagnation point line of the incoming hot gas flow.
  • the hot gas flow is divided into a dividing to the pressure side and a proportion flowing to the suction side deflected to both sides, so that due to the two-sided arrangement of the cooling holes, the underlying component wall is particularly efficiently protected from the high temperatures of the hot gas.
  • the airfoil is part of a turbine blade, in particular a turbine guide vane of a stationary gas turbine.
  • FIG. 1 is a perspective view of a turbine blade 10 is shown.
  • the turbine blade 10 comprises in succession a substantially fir-tree-shaped blade root 12, to which a hot gas platform 14 adjoins.
  • an inventive blade 16 is arranged according to a first embodiment.
  • the airfoil 16 is known to include a leading edge 18 and a trailing edge 20, between which a suction sidewall 17 and a pressure sidewall 19 extend. 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 R 1 , R 2 of cooling holes 22 are arranged.
  • the two rows R 1 , R 2 extend along a wavy line with a plurality of wave troughs and wave crests and are simultaneously arranged on both sides of a middle stagnation point line 24.
  • FIG. 2 A second embodiment of the invention is in FIG. 2 shown.
  • a region is rectilinear, followed by a bulbous section.
  • the two rows R 1 , R 2 of cooling holes 22 in the first, radially inner region are arranged so that they are arranged parallel to the front edge 18 on both sides thereof.
  • This first region B 1 extends between 0% and about 40% vane height.
  • a second region B 2 is provided. This ends at a blade height of about 75%.
  • the cooling holes 22 of both rows R 1 , R 2 move 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 R 1 , R 2 shift back in the direction of the front edge 18.
  • cooling holes 22 are shown only schematically as circles, wherein the throttle cross sections have been shown schematically by different sized circles.
  • the cooling holes 22 may be film cooling holes having a diffuser-like opening. Their diffuser can even be profiled designed.
  • a distance A between the cooling holes 22 to be detected transversely on the surface of the blade 16 may also be different at different blade blade heights.
  • FIG. 3 also shows as a blade profile 28 according to the cross section through the airfoil 16 of the first embodiment FIG. 1 , Between the suction side wall 17 and the pressure side wall 19 extends centrally an imaginary line, which is known as Schaufelprofilmittenline or as Camberline.
  • the blade profile center line is designated by the reference numeral 30.
  • the vorderst arranged point of the blade profile center line 30 defines the leading edge 18.
  • the stagnation point 25 may be slightly shifted away from the leading edge 18 toward the pressure side 19 or towards the suction side 17.
  • the (middle) stagnation points 25 of each blade profile section which can be determined on any blade height, together form the stagnation point line 24.
  • the nose radius is denoted by R.
  • the invention relates to an airfoil 16 for a turbine blade 10, comprising a front edge 18, which can be aspirated by a hot gas S, from which a suction side wall 17 and a pressure side wall 19 extend to a trailing edge 20 of the airfoil 16, the airfoil 16 in a transverse direction thereto extending 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 R 1 , R 2 of cooling holes 22 to one another perpendicular to the front edge 18 to be detected have first distance A.
  • the two rows R 1 , R 2 of cooling holes 22 are at least partially disposed along the front edge 18 on a wavy line.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP18170731.6A 2018-05-04 2018-05-04 Pale d'aube pour une aube de turbine Withdrawn EP3564483A1 (fr)

Priority Applications (7)

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
US17/048,582 US11326458B2 (en) 2018-05-04 2019-05-03 Aerofoil for a turbine blade
PCT/EP2019/061354 WO2019211427A1 (fr) 2018-05-04 2019-05-03 Pale d'aube pour une aube de turbine
KR1020207034682A KR102505046B1 (ko) 2018-05-04 2019-05-03 터빈 블레이드용 에어포일
CN201980030091.4A CN112074652B (zh) 2018-05-04 2019-05-03 用于涡轮叶片的叶片
EP19723730.8A EP3762587B1 (fr) 2018-05-04 2019-05-03 Pale d'aube pour une aube de turbine
JP2020561773A JP7124122B2 (ja) 2018-05-04 2019-05-03 タービン翼用翼形部

Applications Claiming Priority (1)

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

Publications (1)

Publication Number Publication Date
EP3564483A1 true EP3564483A1 (fr) 2019-11-06

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 After (1)

Application Number Title Priority Date Filing Date
EP19723730.8A Active EP3762587B1 (fr) 2018-05-04 2019-05-03 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 (11)

* 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
US11428159B1 (en) 2021-07-01 2022-08-30 Doosan Enerbility Co., Ltd. Airfoil profile for a turbine blade
US11480056B1 (en) 2021-07-01 2022-10-25 Doosan Heavy Industries & Construction Co., Ltd. Airfoil profile for a turbine blade
US11326460B1 (en) 2021-07-15 2022-05-10 Doosan Heavy Industries & Construction Co., Ltd. Airfoil profile for a turbine nozzle
US11454119B1 (en) 2021-07-16 2022-09-27 Doosan Enerbility Co., Ltd Internal core profile for a turbine nozzle airfoil
US11591912B2 (en) 2021-07-16 2023-02-28 Dosan Enerbility Co., Ltd. Internal core profile for a turbine nozzle airfoil
US11377961B1 (en) 2021-07-16 2022-07-05 Doosan Heavy Industries & Construction Co., Ltd. Internal core profile for a turbine nozzle airfoil
US11454125B1 (en) 2021-07-19 2022-09-27 Doosan Heavy Industries & Construction Co., Ltd. Airfoil with directional diffusion region
KR102507408B1 (ko) 2022-11-11 2023-03-08 터보파워텍(주) 3d프린팅에 의한 가스터빈 블레이드의 에어포일 수리 공정
KR20250007369A (ko) 2023-07-05 2025-01-14 강민조 3d프린팅에 의한 가스터빈 블레이드 제작방법

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2154333A2 (fr) 2008-08-14 2010-02-17 United Technologies Corporation Aube refroidie et ensemble de turbine associé
US20160010463A1 (en) * 2013-03-04 2016-01-14 United Technologies Corporation Gas turbine engine high lift airfoil cooling in stagnation zone
EP3043026A2 (fr) * 2014-12-23 2016-07-13 United Technologies Corporation Composant de turbine à gaz, composant d'aube et procédé associé de vectorisation d'air de refroidissement

Family Cites Families (8)

* 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
EP2895696A1 (fr) 2012-08-06 2015-07-22 General Electric Company Composant de turbine rotatif à alignement de trou préférentiel
US10329923B2 (en) * 2014-03-10 2019-06-25 United Technologies Corporation Gas turbine engine airfoil leading edge cooling
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

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2154333A2 (fr) 2008-08-14 2010-02-17 United Technologies Corporation Aube refroidie et ensemble de turbine associé
US20160010463A1 (en) * 2013-03-04 2016-01-14 United Technologies Corporation Gas turbine engine high lift airfoil cooling in stagnation zone
EP3043026A2 (fr) * 2014-12-23 2016-07-13 United Technologies Corporation Composant de turbine à gaz, composant d'aube et procédé associé de vectorisation d'air de refroidissement

Also Published As

Publication number Publication date
JP7124122B2 (ja) 2022-08-23
EP3762587A1 (fr) 2021-01-13
KR102505046B1 (ko) 2023-03-06
JP2021522444A (ja) 2021-08-30
CN112074652A (zh) 2020-12-11
CN112074652B (zh) 2023-05-02
US20210156263A1 (en) 2021-05-27
KR20210002709A (ko) 2021-01-08
WO2019211427A1 (fr) 2019-11-07
EP3762587B1 (fr) 2022-04-13
US11326458B2 (en) 2022-05-10

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