EP0798448A2 - Système et dispositif pour réfroidir une paroi chauffée d'un cÔté par un gaz chaud - Google Patents

Système et dispositif pour réfroidir une paroi chauffée d'un cÔté par un gaz chaud Download PDF

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Publication number
EP0798448A2
EP0798448A2 EP97810115A EP97810115A EP0798448A2 EP 0798448 A2 EP0798448 A2 EP 0798448A2 EP 97810115 A EP97810115 A EP 97810115A EP 97810115 A EP97810115 A EP 97810115A EP 0798448 A2 EP0798448 A2 EP 0798448A2
Authority
EP
European Patent Office
Prior art keywords
cooling
recesses
wall
insert
hot gas
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
EP97810115A
Other languages
German (de)
English (en)
Other versions
EP0798448B1 (fr
EP0798448A3 (fr
Inventor
Günter Dr. Wilfert
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.)
General Electric Switzerland GmbH
Original Assignee
ABB Research Ltd Switzerland
ABB Research Ltd Sweden
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 ABB Research Ltd Switzerland, ABB Research Ltd Sweden filed Critical ABB Research Ltd Switzerland
Publication of EP0798448A2 publication Critical patent/EP0798448A2/fr
Publication of EP0798448A3 publication Critical patent/EP0798448A3/fr
Application granted granted Critical
Publication of EP0798448B1 publication Critical patent/EP0798448B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • 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
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence

Definitions

  • the invention relates to a device and a method for cooling a wall surrounded on one side by hot gas, in particular the hollow profile body of a gas turbine blade, in accordance with the preamble of claim 1.
  • the blown-out cooling air must be redirected as quickly as possible and flow protectively along the profile surface.
  • a rapid lateral expansion of the cooling air is also required.
  • the curvature of the cooling air jets when they exit the holes cause a so-called kidney vortex, i.e. a pair of vertebrae consisting of a right-hand and a left-hand vertebra is generated.
  • this kidney vortex transports part of the hot gas between the holes directly onto the profile surface of the turbine blades and thus under the cooling air jets, which proves to be a serious disadvantage.
  • a major disadvantage of this solution is the weak intensity of the inner vertebra, so that it dissolves relatively quickly and cannot be used permanently to improve the cooling effectiveness.
  • fan-shaped bores are known to improve film cooling.
  • the blowout pulse of the cooling air jet is reduced by means of a diffuser formed in the bore.
  • a faster lateral spread of the cooling air jet is achieved, or an improved film cooling is achieved.
  • the production of such fan-shaped bores is very complex and a wall equipped with such bores is correspondingly expensive.
  • the invention tries to avoid all of these disadvantages. It is based on the task of creating a simple device with an improved cooling effect and a corresponding method for cooling a wall surrounded on one side by hot gas.
  • a radial rib is arranged on the inner surface of the wall upstream of each row of recesses.
  • the cooling insert in the region of the recesses, is deformed in the direction of the wall and is at least approximately parallel to the entry angle of the recesses.
  • the ribs also improve the convective cooling of the wall between the adjacent rows of recesses.
  • the deformation of the cooling insert in the region of the recesses, in the direction of the wall produces both an increased speed of the cooling fluid not flowing into the recesses but further downstream between the wall and the cooling insert, and also a flow directed towards the wall. Due to this additional impingement cooling and the increased flow rate, an improved heat transfer from the wall to the cooling fluid is achieved.
  • a wall cooled in this way can advantageously also be used as a combustion chamber wall or as a heat accumulation segment of a gas turbine.
  • the ribs are arranged up to approximately three times the diameter of the respective recesses, from the center of their entry and protrude approximately half to one diameter of the recesses into the cooling cavity.
  • the cooling insert closes the cooling cavity in the area of the recesses up to a maximum of 30% of the normal distance from the wall and cooling insert.
  • At least one spacer and / or at least one pin are arranged downstream of the recesses in the cooling cavity and connected to the inner surface of the wall.
  • the spacers extend to the cooling insert, while the pins end earlier.
  • the spacers already known from the prior art can be used very effectively in combination with the film cooling according to the invention. Their arrangement exactly between two rows of adjacent recesses is particularly advantageous. In this area, in which almost no film cooling is achieved up to approx. 5 recess diameters downstream of the center of the recess, the spacers act as additional heat sinks for the wall to be cooled, ie they ensure heat flow from the wall to the cooling fluid.
  • pins their additional surface area and the turbulent mixing of the cooling fluid generated with them also act as a heat sink.
  • the guide vane 1 of a gas turbine consists of a hollow profile body 2 which has a wall 3 designed as an outer jacket, a cooling insert 4 arranged at a distance therefrom and a cooling cavity 5 formed between the two.
  • a blade cavity 6 is formed, which is connected in a conventional manner to the compressor of the gas turbine system, not shown, and is acted upon by this with cooling air serving as cooling fluid 7.
  • the outer jacket 3 has an outer and an inner surface 8, 9, between which a plurality of rows of recesses 10 designed as cooling bores are arranged.
  • the blade cavity 6 is connected to the cooling cavity 5 via a plurality of openings 11 arranged in the cooling insert 4 (FIG. 1).
  • the guide vane 1 can also have only a single row of cooling bores 10.
  • hot gas 12 flows out of the combustion chamber (not shown) via the guide blades 1 and the rotor blades of the gas turbine, which are also not shown. Therefore, they have to be constantly cooled.
  • the cooling of the guide vanes 1 takes place by means of the cooling air 7 brought in by the compressor, which penetrates into the cooling cavity 5 via the openings 11 of the cooling insert 4 and initially convectively cools the inner surface 9 of the outer casing 3.
  • the cooling air 7 is then blown out through the cooling bores 10 in a plurality of cooling air jets on the outer surface 8 of the outer casing 3.
  • the curvature of these cooling air jets when they exit into the main flow of hot gas 12 occurs at an exit angle 13 of approximately 30 °.
  • FIG 3 shows an enlarged section of a guide vane 1 designed according to the invention.
  • a streamlined radial rib 15 is arranged on the inner surface 9 of the outer jacket 3 upstream of each row of cooling bores 10.
  • the cooling insert 4 is deformed in the area of the cooling bores 10 in the direction of the outer casing 3 and at least approximately in the process formed parallel to the entry angle 16 of the cooling air 7 in the cooling bores 10.
  • the rib 15 is arranged three times the diameter 17 of the cooling bore 10 away from its entry center 18. At a distance from the outer jacket 3 to the cooling insert 4, which corresponds to twice the diameter 17 of the cooling bore 10, the rib 15 projects a diameter 17 of the cooling bore 10 into the cooling cavity 5. In the area of the cooling bore 10, the cooling insert 4 is deformed in the direction of the outer shell 3 in such a way that it closes the cooling cavity 5 there up to 30% of its normal size.
  • the cooling air 7 is already in the cooling cavity 5, i.e. deflected in the direction upstream of the respective cooling bores 10 in their direction, thus preventing recirculation areas in the cooling bore 10.
  • the center of rotation of this so-called inner vertebrae 19 is not in the center of the cooling bore 10, but in the lower region of the cooling air jet (FIG. 4).
  • the design of the rib 15 leads to a substantially greater deflection of the cooling air 7 when it enters the cooling bores 10.
  • a deflection of approximately 30 ° was customary here, while the cooling air 7 is now deflected at an angle of up to 50 °.
  • the increased deflection of the cooling air 7 and the prevention of a recirculation area in the cooling bores 10 result in a significantly more stable inner vortex 19.
  • this inner vortex 19 is retained even when it exits from each of the cooling bores 10, while the undesired kidney vortex 14 is quickly dissolved in the upper region of the cooling air jet.
  • the inner vortex 19 now ensures that the hot gas 12 is cooled is guided to the outer surface 8 of the outer shell 3 of the guide vane 1.
  • a spacer 20 and a pin 21 are arranged in the cooling cavity 5, downstream of the cooling bore 10 and approximately centrally between two adjacent cooling bores 10. Both the spacer 20 and the pin 21 are connected to the inner surface 9 of the outer jacket 3, the spacer 20 extending to the cooling insert 4 and the pin 21 being shorter. Due to the central arrangement of spacer 20 and pin 21 between two adjacent cooling bores 10, the area with the lowest cooling effect, i.e. up to about five diameters 17 downstream of the outlet center point 22 of the cooling bore, sufficient heat flow from the outer jacket 3 to the cooling air 7 is achieved.
  • Such a cooling configuration is of course not restricted to the guide blades 1 of gas turbines. It can also be used in rotor blades, combustion chamber walls, heat accumulation segments of gas turbines or in other walls 3 surrounded on one side by hot gas 12.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Cooling Or The Like Of Electrical Apparatus (AREA)
EP97810115A 1996-03-30 1997-03-03 Système et dispositif pour réfroidir une paroi chauffée d'un côté par un gaz chaud Expired - Lifetime EP0798448B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE19612840A DE19612840A1 (de) 1996-03-30 1996-03-30 Vorrichtung und Verfahren zur Kühlung einer einseitig von Heissgas umgebenen Wand
DE19612840 1996-03-30

Publications (3)

Publication Number Publication Date
EP0798448A2 true EP0798448A2 (fr) 1997-10-01
EP0798448A3 EP0798448A3 (fr) 1999-05-06
EP0798448B1 EP0798448B1 (fr) 2003-05-07

Family

ID=7790044

Family Applications (1)

Application Number Title Priority Date Filing Date
EP97810115A Expired - Lifetime EP0798448B1 (fr) 1996-03-30 1997-03-03 Système et dispositif pour réfroidir une paroi chauffée d'un côté par un gaz chaud

Country Status (4)

Country Link
US (1) US5779438A (fr)
EP (1) EP0798448B1 (fr)
JP (1) JP3886593B2 (fr)
DE (2) DE19612840A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2233693A1 (fr) * 2008-01-08 2010-09-29 IHI Corporation Structure de refroidissement d'aube de turbine
WO2013123115A1 (fr) 2012-02-15 2013-08-22 United Technologies Corporation Composant de moteur à turbine à gaz comportant un refroidissement par contact et diffusion
EP2815078A4 (fr) * 2012-02-15 2015-12-30 United Technologies Corp Composant de moteur à turbine à gaz ayant un trou de contact et de refroidissement à lobes
WO2016178689A1 (fr) * 2015-05-07 2016-11-10 Siemens Aktiengesellschaft Aube de turbine avec canaux de refroidissement internes
EP3176374A1 (fr) * 2015-12-03 2017-06-07 General Electric Company Refroidissement de bord de fuite pour une pale de turbine

Families Citing this family (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6474947B1 (en) * 1998-03-13 2002-11-05 Mitsubishi Heavy Industries, Ltd. Film cooling hole construction in gas turbine moving-vanes
DE19845147B4 (de) 1998-10-01 2006-11-23 Alstom Vorrichtung und Verfahren zur Kühlung einer einseitug von Heißgas umgebenen Wand
GB2350867B (en) * 1999-06-09 2003-03-19 Rolls Royce Plc Gas turbine airfoil internal air system
US7128533B2 (en) * 2004-09-10 2006-10-31 Siemens Power Generation, Inc. Vortex cooling system for a turbine blade
FR2890103A1 (fr) * 2005-08-25 2007-03-02 Snecma Deflecteur d'air pour circuit de refroidissement pour aube de turbine a gaz
JP4147239B2 (ja) 2005-11-17 2008-09-10 川崎重工業株式会社 ダブルジェット式フィルム冷却構造
US8109724B2 (en) * 2009-03-26 2012-02-07 United Technologies Corporation Recessed metering standoffs for airfoil baffle
US9328616B2 (en) * 2013-02-01 2016-05-03 Siemens Aktiengesellschaft Film-cooled turbine blade for a turbomachine
US10655855B2 (en) 2013-08-30 2020-05-19 Raytheon Technologies Corporation Gas turbine engine wall assembly with support shell contour regions
US9708915B2 (en) 2014-01-30 2017-07-18 General Electric Company Hot gas components with compound angled cooling features and methods of manufacture
US10533745B2 (en) 2014-02-03 2020-01-14 United Technologies Corporation Film cooling a combustor wall of a turbine engine
WO2015184294A1 (fr) 2014-05-29 2015-12-03 General Electric Company Générateur de turbulence fastback
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US10260353B2 (en) * 2014-12-04 2019-04-16 Rolls-Royce Corporation Controlling exit side geometry of formed holes
CN104594956B (zh) * 2015-02-10 2016-02-17 河北工业大学 一种提高开槽气膜孔下游壁面气膜冷却效率的结构
CN105401983B (zh) * 2015-12-24 2017-04-12 河北工业大学 一种提高组件外部冷却效果的上游结构
US10648341B2 (en) 2016-11-15 2020-05-12 Rolls-Royce Corporation Airfoil leading edge impingement cooling
US10465526B2 (en) 2016-11-15 2019-11-05 Rolls-Royce Corporation Dual-wall airfoil with leading edge cooling slot
US10450873B2 (en) 2017-07-31 2019-10-22 Rolls-Royce Corporation Airfoil edge cooling channels
US11408302B2 (en) * 2017-10-13 2022-08-09 Raytheon Technologies Corporation Film cooling hole arrangement for gas turbine engine component
US10570751B2 (en) 2017-11-22 2020-02-25 General Electric Company Turbine engine airfoil assembly
GB201806821D0 (en) * 2018-04-26 2018-06-13 Rolls Royce Plc Coolant channel
CN113217462B (zh) * 2021-06-08 2022-11-29 西北工业大学 亚声速旋涡吹气式压气机叶片

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US3628885A (en) * 1969-10-01 1971-12-21 Gen Electric Fluid-cooled airfoil
US4153386A (en) * 1974-12-11 1979-05-08 United Technologies Corporation Air cooled turbine vanes
JPS55104507A (en) * 1979-02-05 1980-08-11 Ishikawajima Harima Heavy Ind Co Ltd Cooling blade for high-temperature turbine
EP0258754A2 (fr) * 1986-09-03 1988-03-09 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Aube de turbine avec noyau de refroidissement
JPH07145702A (ja) * 1993-11-22 1995-06-06 Toshiba Corp タービン冷却翼

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JPH0352504A (ja) * 1989-07-18 1991-03-06 Fujikura Ltd 管路内自走装置

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Publication number Priority date Publication date Assignee Title
US3628885A (en) * 1969-10-01 1971-12-21 Gen Electric Fluid-cooled airfoil
US4153386A (en) * 1974-12-11 1979-05-08 United Technologies Corporation Air cooled turbine vanes
JPS55104507A (en) * 1979-02-05 1980-08-11 Ishikawajima Harima Heavy Ind Co Ltd Cooling blade for high-temperature turbine
EP0258754A2 (fr) * 1986-09-03 1988-03-09 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Aube de turbine avec noyau de refroidissement
JPH07145702A (ja) * 1993-11-22 1995-06-06 Toshiba Corp タービン冷却翼

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PATENT ABSTRACTS OF JAPAN vol. 004, no. 154 (M-038), 28. Oktober 1980 & JP 55 104507 A (ISHIKAWAJIMA HARIMA HEAVY IND CO LTD), 11. August 1980 *
PATENT ABSTRACTS OF JAPAN vol. 095, no. 009, 31. Oktober 1995 & JP 07 145702 A (TOSHIBA CORP), 6. Juni 1995 & US 5 533 864 A (NOMOTO ET AL) 9. Juli 1996 *
WILFERT G]NTER: "Experimentelle und numerische Untersuchungen der Mischungsvorgange zwischen Kuhlfilmen und Gitterstromung an einem hobhbelasteten Turbinengitter" AERODYNAMIK DER FILMK]HLUNG (ABSCHLUSSBERICHT, VORHABEN NR. 520 UND 594), Bd. 573, 1. Juli 1991 - 31. Dezember 1994, XP002095932 *

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2233693A1 (fr) * 2008-01-08 2010-09-29 IHI Corporation Structure de refroidissement d'aube de turbine
EP2233693A4 (fr) * 2008-01-08 2011-03-16 Ihi Corp Structure de refroidissement d'aube de turbine
US9133717B2 (en) 2008-01-08 2015-09-15 Ihi Corporation Cooling structure of turbine airfoil
WO2013123115A1 (fr) 2012-02-15 2013-08-22 United Technologies Corporation Composant de moteur à turbine à gaz comportant un refroidissement par contact et diffusion
EP2815079A4 (fr) * 2012-02-15 2015-12-30 United Technologies Corp Composant de moteur à turbine à gaz comportant un refroidissement par contact et diffusion
EP2815078A4 (fr) * 2012-02-15 2015-12-30 United Technologies Corp Composant de moteur à turbine à gaz ayant un trou de contact et de refroidissement à lobes
WO2016178689A1 (fr) * 2015-05-07 2016-11-10 Siemens Aktiengesellschaft Aube de turbine avec canaux de refroidissement internes
EP3176374A1 (fr) * 2015-12-03 2017-06-07 General Electric Company Refroidissement de bord de fuite pour une pale de turbine
US10344598B2 (en) 2015-12-03 2019-07-09 General Electric Company Trailing edge cooling for a turbine blade
US11208901B2 (en) 2015-12-03 2021-12-28 General Electric Company Trailing edge cooling for a turbine blade

Also Published As

Publication number Publication date
JP3886593B2 (ja) 2007-02-28
EP0798448B1 (fr) 2003-05-07
DE19612840A1 (de) 1997-10-02
US5779438A (en) 1998-07-14
DE59710009D1 (de) 2003-06-12
EP0798448A3 (fr) 1999-05-06
JPH108909A (ja) 1998-01-13

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