WO2008059620A1 - Structure de refroidissement par film - Google Patents

Structure de refroidissement par film Download PDF

Info

Publication number
WO2008059620A1
WO2008059620A1 PCT/JP2007/054910 JP2007054910W WO2008059620A1 WO 2008059620 A1 WO2008059620 A1 WO 2008059620A1 JP 2007054910 W JP2007054910 W JP 2007054910W WO 2008059620 A1 WO2008059620 A1 WO 2008059620A1
Authority
WO
WIPO (PCT)
Prior art keywords
film cooling
combustion gas
hole
cooling structure
film
Prior art date
Application number
PCT/JP2007/054910
Other languages
English (en)
Japanese (ja)
Inventor
Yoji Ohkita
Original Assignee
Ihi Corporation
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 Ihi Corporation filed Critical Ihi Corporation
Priority to US12/514,511 priority Critical patent/US20100040459A1/en
Priority to EP07738382.6A priority patent/EP2083147B1/fr
Priority to CA2668750A priority patent/CA2668750C/fr
Publication of WO2008059620A1 publication Critical patent/WO2008059620A1/fr

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
    • F05D2250/00—Geometry
    • F05D2250/50—Inlet or outlet
    • F05D2250/52—Outlet
    • 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 present invention relates to a film cooling structure suitable for film cooling a surface of a component (turbine blade or the like) of a gas turbine engine.
  • a gas turbine engine increases with increasing combustion gas temperature.
  • This combustion gas heats the structural walls of components (combustor liners, turbine blades, turbine shrouds, etc.) placed in the combustion gas flow path. Heat to. Therefore, in order to efficiently cool the structural walls of such components, a cooling passage is provided in the interior, and cooling air is flowed to perform convection cooling, and cooling air is exposed to high-temperature combustion gas from the film cooling holes.
  • a film cooling structure is employed in which film cooling is performed by ejecting a film on the surface (see, for example, Patent Documents:! To 5 below).
  • FIG. 1A to FIG. 1C show an example of a conventional film cooling structure 30.
  • Figure 1B is 1B of Figure 1A
  • FIG. 1B is a sectional view taken along line B
  • FIG. 1C is a sectional view taken along line 1C-1C in FIG. 1B.
  • the structural wall 31 has a surface 32 that is exposed to the combustion gas 1 and an inner surface 33 that is located on the opposite side of the surface 32.
  • film cooling holes 34 for guiding the cooling air 5 on the inner surface 33 side to the surface 32 side and cooling the film on the surface 32 are formed at a predetermined angle with respect to the surface 32.
  • the film cooling hole 34 has an introduction portion 34a extending from the inner surface 33 toward the surface 32 to a middle position in the structural wall 31, and a cross-sectional area gradually increasing from the surface 32 side end portion of the introduction portion 34a toward the surface 32. It has an enlarged portion 34b (diffuser) that opens at the surface 32. As shown in FIG.
  • Patent Document 1 Japanese Unexamined Patent Application Publication No. 2006-9785
  • Patent Document 2 JP 2005-90511
  • Patent Document 3 Japanese Patent Laid-Open No. 2003-41902
  • Patent Document 4 Japanese Patent Laid-Open No. 2001-173405
  • Patent Document 5 Japanese Patent Laid-Open No. 10-89005
  • the present invention has been made in view of such problems, and provides a film cooling structure that can increase the expansion angle at the expansion portion and can improve the average film cooling efficiency. For the purpose.
  • the film cooling structure of the present invention employs the following means.
  • the present invention comprises a structural wall having a surface exposed to combustion gas and an inner surface located on the opposite side of the surface, and the cooling air on the inner surface side is guided to the structural wall to the surface side for film cooling of the surface.
  • the film cooling holes include an introduction part extending from the inner surface toward the surface to a middle position in the structure wall, and a surface side of the introduction part.
  • An enlarged portion whose cross-sectional area gradually increases from the end toward the surface and opens at the surface, and a partition portion that divides the inside of the enlarged portion in a hole width direction perpendicular to the flow direction of the combustion gas. It is characterized by having.
  • the film cooling hole has the partition portion configured as described above, the effective area enlargement rate can be suppressed. Therefore, even if the lateral enlargement angle of the enlargement portion is increased, the cooling air can be reduced. Qi peeling is suppressed. Therefore, the cooling air can be effectively diffused compared to the conventional technology, and the enlargement angle in the lateral direction of the enlargement portion can be increased. Cooling air can be spread thinly and widely on the surface of the structural wall to improve average film cooling efficiency. The definition of the average film cooling efficiency will be described later.
  • the cooling air can be spread more thinly and widely on the surface of the structural wall as compared with the prior art, the number of film cooling holes formed in the structural wall can be reduced. For this reason, the manufacturing process of the film cooling structure can be reduced. Further, since the amount of cooling air extracted from the compressor of the gas turbine engine can be reduced as the number of film cooling holes is reduced, the engine efficiency can be improved.
  • the partition portion is formed at an intermediate position in the hole width direction perpendicular to the flow direction of the combustion gas inside the film cooling hole, and the flow direction of the combustion gas One of the wall surface facing the upstream side and the wall surface facing the downstream side.
  • the force protrudes toward the other side and extends from the inner surface of the structural wall toward the surface over the entire area inside the hole.
  • the partition portion does not completely partition the film cooling hole in the lateral direction and extends over the entire region in the thickness direction of the structural wall, so that the film cooling hole can be easily processed.
  • FIG. 1A is a plan view showing a conventional film cooling structure.
  • FIG. 1B is a cross-sectional view taken along line 1B-1B in FIG. 1A.
  • 1C is a cross-sectional view taken along line 1C-1C in FIG. 1B.
  • FIG. 2 is a perspective view of a turbine rotor blade to which the film cooling structure of the present invention is applied.
  • FIG. 3A is a plan view showing a film cooling structure that applies force to an embodiment of the present invention.
  • 3B is a cross-sectional view taken along line 3B-3B in FIG. 3A.
  • FIG. 3C is a cross-sectional view taken along line 3C-3C in FIG. 3B.
  • FIG. 4 is a perspective view showing the shape of a film cooling hole in a film cooling structure that works according to an embodiment of the present invention.
  • FIG. 5 is a diagram for explaining the physical action of the partition part.
  • the film cooling structure of the present invention is applied to components arranged in a combustion gas flow path in a gas turbine engine. These components include a combustor liner, a turbine nose vane, a turbine nose nore band, a turbine blade, a turbine stationary blade, a turbine shroud, and a turbine exhaust liner.
  • FIG. 2 shows a perspective view of a turbine rotor blade 2 to which the film cooling structure 10 of the present invention is applied.
  • the turbine rotor blade 2 includes a blade portion 3 as a structural wall having a surface 12 exposed to the combustion gas 1 and a base portion 4 for mounting the blade portion 3 on an engine rotor.
  • a cooling circuit (not shown) for flowing cooling air is formed inside the wing part 3. This cooling air is extracted from the compressor of the gas turbine engine and flows into the cooling circuit via a flow path (not shown) formed in the base portion 4. Cooling air that has flowed into the cooling circuit is ejected from a large number of film cooling holes 14 provided on the surface 12 of the wing 3 to cool the surface 12 of the wing 3.
  • FIGS. 3A-3C illustrate a film cooling structure 10 that works with embodiments of the present invention.
  • FIG. 3A is a plan view showing the film cooling structure 10.
  • 3B is a cross-sectional view taken along line 3B-3B of FIG. 3A.
  • 3C is a cross-sectional view taken along line 3C-3C of FIG. 3B.
  • FIG. 4 is a perspective view showing the shape of the film cooling hole 14 in the film cooling structure 10 according to the embodiment of the present invention.
  • the film cooling structure 10 is applied to components such as turbine rotor blades arranged in the flow path of the combustion gas 1 in a gas turbine engine.
  • the film cooling structure 10 includes a structural wall 11 having a surface 12 exposed to the combustion gas 1 and an inner surface 13 located on the opposite side of the surface 12.
  • the wall constituting the blade portion of the turbine rotor blade is the structural wall 11. Cooling air 5 flows on the inner surface 13 side of the structural wall 11.
  • film cooling holes 14 for guiding the cooling air 5 on the inner surface 13 side to the surface 12 side and cooling the film on the surface 12 are formed.
  • the axis of the film cooling hole 14 is inclined at a predetermined angle with respect to the surface 12 of the structural wall 11 so that the cooling air 5 is blown out in the direction along the flow of the combustion gas 1. is doing.
  • the film cooling hole 14 has an introduction portion 14a extending from the inner surface 13 toward the surface 12 to an intermediate position in the structural wall 11, and a cross-sectional area from the end of the introduction portion 14a on the surface 12 side toward the surface 12. Has an enlarged portion 14b that gradually increases and opens at the surface 12.
  • the film cooling hole 14 further includes a partitioning portion 16 that partitions the inside of the enlarged portion 14b into a plurality of holes in a hole width direction orthogonal to the flow direction of the combustion gas 1.
  • the “hole width direction perpendicular to the flow direction of the combustion gas 1” is a direction perpendicular to the paper surface in FIG. 3B, and a horizontal direction in FIG. 3C.
  • the partition 16 is formed at an intermediate position in the hole width direction perpendicular to the flow direction of the combustion gas 1 inside the film cooling hole 14. Projecting from the wall surface facing the upstream side in the flow direction toward the upstream side in the flow direction of the combustion gas 1 and extending from the inner surface 13 to the surface 12 of the structural wall 11 over the entire interior of the hole. A gap is formed between the partition portion 16 and the wall surface facing the downstream side in the flow direction of the combustion gas 1.
  • a single partitioning portion 16 may be provided at intervals in the hole width direction.
  • the partition 16 is provided so as to protrude from the wall surface facing the upstream side in the flow direction of the combustion gas 1 toward the upstream side in the flow direction of the combustion gas 1.
  • it may be provided so as to protrude from the wall surface facing the downstream side in the flow direction of the combustion gas 1 toward the downstream side in the flow direction of the combustion gas 1.
  • a gap is formed between the partition portion 16 and the wall surface facing the upstream side in the flow direction of the combustion gas 1.
  • Fig. 5 shows a graph of the diffuser with the length ratio on the logarithmic scale on the horizontal axis, the inlet / outlet area ratio minus 1 on the logarithmic scale on the vertical axis, and the pressure recovery rate (deceleration rate) Cp as a parameter. Show. At this time, in the case of the same inlet / outlet area ratio, the larger the length ratio, the smaller the enlargement angle. Further, peeling is less likely to occur when the pressure recovery rate is higher.
  • the straight line indicated by the pressure recovery rate Cp * * in the figure is a line connecting the points where the maximum pressure recovery rate is obtained when the diffuser inlet / outlet area ratio is constant.
  • the Cp * straight line is the line that provides the maximum pressure recovery rate when the length ratio is constant. Therefore, if the inlet / outlet area ratio is constant, It can be seen that the smaller the angle of enlargement, the higher the pressure recovery rate and the less likely the peeling occurs. Dividing the diffuser's passage into two or three equal parts, the expansion angle of each small passage becomes 1/2 or 1/3 of the overall expansion angle, and is smaller than the expansion angle determined by Cp *. As a result, a high pressure recovery rate can be obtained.
  • the film cooling hole 14 includes the cutting portion 16 configured as described above, the effective area enlargement ratio can be suppressed, and thus the enlargement of the enlargement portion 14b in the lateral direction. Even if the angle is increased, the separation of the cooling air 5 is suppressed. For this reason, since the cooling air 5 can be effectively diffused as compared with the prior art, the lateral expansion angle of the enlarged portion 14b can be increased. It can be spread thinly and widely to improve the average film cooling efficiency.
  • the average film cooling efficiency is given by (fuel gas temperature / structure wall surface temperature) / (combustion gas temperature / cooling air temperature).
  • the cooling air 5 can be spread more thinly and widely on the surface 12 of the structural wall 11 as compared with the prior art, the number of film cooling holes 14 formed in the structural wall 11 can be reduced. For this reason, the manufacturing process of the film cooling structure 10 can be reduced. Further, as the number of film cooling holes 14 decreases, the amount of cooling air extracted from the compressor of the gas turbine engine can be reduced, so that the engine efficiency can be improved.
  • the partition 16 When the film cooling holes 14 are processed by a method such as a discharge casing, if the partition 16 completely partitions the film cooling holes 14 in the lateral direction, the discharge casing is divided for each of the divided holes. It is necessary to insert the electrode and clean the hole. Further, if the partition 16 has a shape that is interrupted at a position in the thickness direction of the structural wall 11, a plurality of steps are required to cover one film cooling hole 14 (for example, the surface 12 side and the inner surface It is necessary to insert and process the discharge casing electrode from the 13th side). Further, the processing steps are similarly complicated in other processing means.
  • the partition 16 does not completely partition the film cooling hole 14 in the lateral direction, and extends over the entire region in the thickness direction of the structural wall 11, so that it is shown in FIGS. 3A to 3C and FIG.
  • a discharge carriage electrode configured to cover the film cooling hole 14 from the surface 12 side
  • the film cooling hole 14 can be covered in a single step. Therefore, the film cooling hole 14 can be easily processed.
  • the present invention is applied to the turbine rotor blade 2, but the combustor liner, the turbine nozle vane, the turbine nozle band, the turbine stationary blade arranged in the flow path of the combustion gas in the gas turbine engine. Also applicable to turbine shrouds, turbine exhaust liners and other components.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gas Burners (AREA)

Abstract

L'invention concerne une structure de refroidissement par film (10) comprenant une paroi de structure (11) présentant une surface externe (12) exposée à un gaz de combustion et une surface interne (13) positionnée en face de la surface externe (12), la structure de refroidissement par film (10) étant dotée d'un orifice de refroidissement par film (14) pour, dans la paroi de structure (11), diriger de l'air de refroidissement sur le côté de surface interne (13) vers le côté de surface externe (12) pour ainsi réaliser un refroidissement par film sur la surface externe (12). L'orifice de refroidissement par film (14) comporte une partie d'introduction (14a) s'étendant depuis la surface interne (13) vers la surface externe (12) jusqu'à une position médiane dans la paroi de structure (11) ; la zone de coupe transversale d'une partie d'agrandissement (14b) augmentant graduellement depuis l'extrémité de la partie d'introduction (14a) sur le côté de surface externe (12) vers la surface externe (12) et s'ouvrant sur la surface externe (12) ; et une partie de séparation (16) pour diviser l'intérieur de la partie d'agrandissement (14b) en des unités multiples dans la direction de la largeur de l'orifice, perpendiculaire à la direction d'écoulement du gaz de combustion.
PCT/JP2007/054910 2006-11-13 2007-03-13 Structure de refroidissement par film WO2008059620A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US12/514,511 US20100040459A1 (en) 2006-11-13 2007-03-13 Film cooling structure
EP07738382.6A EP2083147B1 (fr) 2006-11-13 2007-03-13 Structure de refroidissement par film
CA2668750A CA2668750C (fr) 2006-11-13 2007-03-13 Structure de refroidissement par film

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2006-306538 2006-11-13
JP2006306538A JP4941891B2 (ja) 2006-11-13 2006-11-13 フィルム冷却構造

Publications (1)

Publication Number Publication Date
WO2008059620A1 true WO2008059620A1 (fr) 2008-05-22

Family

ID=39401434

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/JP2007/054910 WO2008059620A1 (fr) 2006-11-13 2007-03-13 Structure de refroidissement par film

Country Status (5)

Country Link
US (1) US20100040459A1 (fr)
EP (1) EP2083147B1 (fr)
JP (1) JP4941891B2 (fr)
CA (1) CA2668750C (fr)
WO (1) WO2008059620A1 (fr)

Cited By (3)

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JP2012052535A (ja) * 2010-08-31 2012-03-15 General Electric Co <Ge> 共形湾曲フィルム孔を備えた構成要素及びその製造方法
WO2013089255A1 (fr) * 2011-12-15 2013-06-20 株式会社Ihi Lame de turbine
US11414999B2 (en) * 2016-07-11 2022-08-16 Raytheon Technologies Corporation Cooling hole with shaped meter

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US8371814B2 (en) * 2009-06-24 2013-02-12 Honeywell International Inc. Turbine engine components
US8529193B2 (en) * 2009-11-25 2013-09-10 Honeywell International Inc. Gas turbine engine components with improved film cooling
US8628293B2 (en) 2010-06-17 2014-01-14 Honeywell International Inc. Gas turbine engine components with cooling hole trenches
US8858175B2 (en) * 2011-11-09 2014-10-14 General Electric Company Film hole trench
US20130209235A1 (en) * 2012-02-15 2013-08-15 United Technologies Corporation Gas turbine engine component with cusped, lobed cooling hole
US8689568B2 (en) 2012-02-15 2014-04-08 United Technologies Corporation Cooling hole with thermo-mechanical fatigue resistance
US9422815B2 (en) * 2012-02-15 2016-08-23 United Technologies Corporation Gas turbine engine component with compound cusp cooling configuration
US8763402B2 (en) * 2012-02-15 2014-07-01 United Technologies Corporation Multi-lobed cooling hole and method of manufacture
US8584470B2 (en) * 2012-02-15 2013-11-19 United Technologies Corporation Tri-lobed cooling hole and method of manufacture
US8683814B2 (en) * 2012-02-15 2014-04-01 United Technologies Corporation Gas turbine engine component with impingement and lobed cooling hole
US9273560B2 (en) * 2012-02-15 2016-03-01 United Technologies Corporation Gas turbine engine component with multi-lobed cooling hole
US8850828B2 (en) 2012-02-15 2014-10-07 United Technologies Corporation Cooling hole with curved metering section
US10422230B2 (en) 2012-02-15 2019-09-24 United Technologies Corporation Cooling hole with curved metering section
US9024226B2 (en) * 2012-02-15 2015-05-05 United Technologies Corporation EDM method for multi-lobed cooling hole
US9284844B2 (en) * 2012-02-15 2016-03-15 United Technologies Corporation Gas turbine engine component with cusped cooling hole
US8683813B2 (en) * 2012-02-15 2014-04-01 United Technologies Corporation Multi-lobed cooling hole and method of manufacture
US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
EP2861909A2 (fr) * 2012-06-13 2015-04-22 General Electric Company Paroi de moteur de turbine à gaz
US10113433B2 (en) * 2012-10-04 2018-10-30 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
CN105189976B (zh) * 2013-03-15 2019-04-16 联合工艺公司 用于在冷却孔内附加特征的加成制造方法
EP2886798B1 (fr) * 2013-12-20 2018-10-24 Rolls-Royce Corporation Trous de refroidissement de film usinés mécaniquement
CN106068371B (zh) * 2014-04-03 2018-06-08 三菱日立电力系统株式会社 叶片分割体、叶片组列、燃气涡轮机
US11313235B2 (en) 2015-03-17 2022-04-26 General Electric Company Engine component with film hole
US10208602B2 (en) * 2015-04-27 2019-02-19 United Technologies Corporation Asymmetric diffuser opening for film cooling holes
US11021965B2 (en) 2016-05-19 2021-06-01 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
KR101853550B1 (ko) 2016-08-22 2018-04-30 두산중공업 주식회사 가스 터빈 블레이드
US20180058770A1 (en) * 2016-09-01 2018-03-01 Additive Rocket Corporation Structural heat exchanger
EP3354849A1 (fr) * 2017-01-31 2018-08-01 Siemens Aktiengesellschaft Paroi pour composant à gaz chaud et composant à gaz chaud associé pour turbine à gaz
WO2019030389A1 (fr) 2017-08-11 2019-02-14 Archroma Ip Gmbh Solutions aqueuses concentrées purifiées de sel de leucoindigo
US10933481B2 (en) * 2018-01-05 2021-03-02 General Electric Company Method of forming cooling passage for turbine component with cap element
US11286792B2 (en) * 2019-07-30 2022-03-29 Rolls-Royce Plc Ceramic matrix composite vane with cooling holes and methods of making the same

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2012052535A (ja) * 2010-08-31 2012-03-15 General Electric Co <Ge> 共形湾曲フィルム孔を備えた構成要素及びその製造方法
WO2013089255A1 (fr) * 2011-12-15 2013-06-20 株式会社Ihi Lame de turbine
US10060265B2 (en) 2011-12-15 2018-08-28 Ihi Corporation Turbine blade
US11414999B2 (en) * 2016-07-11 2022-08-16 Raytheon Technologies Corporation Cooling hole with shaped meter

Also Published As

Publication number Publication date
EP2083147A1 (fr) 2009-07-29
JP4941891B2 (ja) 2012-05-30
US20100040459A1 (en) 2010-02-18
JP2008121561A (ja) 2008-05-29
EP2083147A4 (fr) 2014-05-14
CA2668750A1 (fr) 2008-05-22
EP2083147B1 (fr) 2015-10-07
CA2668750C (fr) 2012-06-19

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