US20140037429A1 - Turbine vane - Google Patents

Turbine vane Download PDF

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Publication number
US20140037429A1
US20140037429A1 US14/046,191 US201314046191A US2014037429A1 US 20140037429 A1 US20140037429 A1 US 20140037429A1 US 201314046191 A US201314046191 A US 201314046191A US 2014037429 A1 US2014037429 A1 US 2014037429A1
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United States
Prior art keywords
vane
hole
turbine
film cooling
cooling
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.)
Abandoned
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US14/046,191
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English (en)
Inventor
Yoji Okita
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.)
IHI Corp
Society of Japanese Aerospace Companies
Original Assignee
IHI Corp
Society of Japanese Aerospace Companies
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 Corp, Society of Japanese Aerospace Companies filed Critical IHI Corp
Assigned to IHI CORPORATION, THE SOCIETY OF JAPANESE AEROSPACE COMPANIES, INC. reassignment IHI CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: OKITA, YOJI
Publication of US20140037429A1 publication Critical patent/US20140037429A1/en
Abandoned legal-status Critical Current

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    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/12—Two-dimensional rectangular
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/14—Two-dimensional elliptical
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a turbine vane used in a turbine for gas turbine engines such as an aircraft engine and an industrial gas turbine engine.
  • turbine vanes to be exposed to a combustion gas while a gas turbine engine is in operation can be cooled by use of cooling air (part of compressed air) extracted from a compressor or a fan of the gas turbine engine.
  • a cooling passage into which the cooling air can flow is formed inside the turbine vane, and multiple film cooling holes through which the cooling air can jet out are formed in a vane surface of the turbine vane in such a way as to communicate with the cooling passage. For this reason, while the gas turbine engine is in operation, the cooling air flowing into the cooling passage is jetted out through the multiple film cooling holes, thus forms a film cooling layer which covers the vane surface of the turbine vane, and can perform film cooling on the turbine vane.
  • an object of the present invention is to provide a turbine vane having a novel configuration which is capable of fully cooling a front edge-side vane surface and its vicinity.
  • An aspect of the present invention is a turbine vane for a turbine of a gas turbine engine, and capable of being cooled by cooling air, the turbine vane comprising: a vane body including: a vane surface; a cooling passage allowing the cooling air to flow into the vane body; and a plurality of film cooling holes formed in the vane surface on a front edge side of the vane body so as to communicate with the cooling passage to jet out the cooling air through the plurality of film cooling holes, a hole cross section of each film cooling hole having a long-hole shape extending in a direction parallel to a cross section along a span direction of the vane body, a hole-center line of each film cooling hole tilting from a thickness direction of the vane body on a cross section of the vane surface along the span direction, and an exit-side and obtuse angle-side portion of a hole wall surface of each film cooling hole tilting further from the thickness direction than the hole-center line on the cross section along the span direction.
  • the “turbine vane” represents a turbine rotor vane and a turbine stator vane
  • the “hole cross section” means a cross section perpendicular to the hole-center line.
  • the “exit side” represents the exit side viewed in the flowing direction of the cooling air
  • the “obtuse angle side” represents a side (region) having an obtuse angle defined with the vane surface.
  • the present invention makes it possible to fully diffuse the cooling air, which is jetted out through each film cooling hole, on the front edge-side vane surface of the turbine vane in the span direction while inhibiting the cooling air from coming up away from the front edge-side vane surface of the turbine vane. For this reason, it is possible to increase the film efficiency around the front edge-side vane surface of the turbine vane, and accordingly to increase the cooling performance of the turbine vane to a higher level.
  • FIG. 1 is a perspective view of a turbine stator vane of an embodiment of the present invention.
  • FIG. 2 is a side view of the turbine stator vane of the embodiment of the present invention, which is partially cut away.
  • FIG. 3 is a magnified cross-sectional view of the turbine stator vane taken along a line in FIG. 2 .
  • FIG. 4A is a cross-sectional view of a part of a front edge-side vane surface of the turbine stator vane of the embodiment of the present invention, which extends in a span direction
  • FIG. 4B is a magnified cross-sectional view taken along a IVB-IVB line in FIG. 4A .
  • FIG. 4C is a magnified view of a portion viewed in a direction indicated with an arrow IVC in FIG. 4A .
  • FIG. 4D is a magnified view of a portion viewed in a direction indicated with an arrow IVD in FIG. 4A .
  • FIG. 5A is a cross-sectional view of a front edge-side portion of a turbine vane of Invention Example.
  • FIG. 5B is a cross-sectional view taken along a VB-VB line in FIG. 5A .
  • FIG. 6A is a cross-sectional view of a front edge-side portion of a turbine vane of Comparative Example 1.
  • FIG. 6B is a cross-sectional view taken along a XIB-VIB line in FIG. 6A .
  • FIG. 7A is a cross-sectional view of a front edge-side portion of a turbine vane of Comparative Example 2.
  • FIG. 7B is a cross-sectional view taken along a XIIB-VIIB line in FIG. 7A .
  • FIG. 8 is a schematic diagram for explaining a stagnation point and the like on the front edge-side vane surface of the turbine vane.
  • FIGS. 9A , 9 B and 9 C are diagrams showing how the temperature is distributed around the front edge-side vane surface of the turbine vane of Invention Example.
  • FIGS. 10A , 10 B and 10 C are diagrams showing how the temperature is distributed around the front edge-side vane surface of the turbine vane of Comparative Example 1.
  • FIGS. 11A , 11 B and 11 C are diagrams showing how the temperature is distributed around the front edge-side vane surface of the turbine vane of Comparative Example 2.
  • FIG. 12 is a diagram showing relationships between the distance from the stagnation point on the front edge-side vane surface and film efficiency regarding the turbine vane of Invention Example, the turbine vane of Comparative Example 1 and the turbine vane of Comparative Example 2.
  • reference sign “FF” denotes a frontward direction
  • reference sign “FR” denotes a rearward direction.
  • the frontward direction means an upstream direction in the light of a flowing direction of a main stream
  • the rearward direction means a downstream direction in the light of the flowing direction of the main stream.
  • a turbine stator vane 1 of the embodiment is used for a turbine (not shown) of a gas turbine engine such as an aircraft engine and an industrial gas turbine engine.
  • the turbine stator vane 1 can be cooled by cooling air (part of compressed air) CA extracted from a compressor (not shown) or a fan (not shown) of the gas turbine engine.
  • the specific configuration of the turbine stator vane 1 is as follows.
  • the turbine stator vane 1 is produced (cast) by lost wax precision casting, for example.
  • the turbine stator vane 1 includes a hollow stator vane body (vane body) 3 .
  • a rib wall (partition wall) 5 is provided inside the stator vane body 3 .
  • the rib wall 5 divides the inside of the stator vane body 3 into front and rear halves, and forms a front cooling passage 7 and a rear cooling passage 9 into which the cooling air CA can flow.
  • the front cooling passage 7 is formed in the front half of the stator vane body 3
  • the rear cooling passage 9 is formed in the rear half of the stator vane body 3 .
  • the stator vane body 3 has a vane surface 3 a in its front edge-side portion.
  • Multiple film cooling holes 11 are formed in the vane surface 3 a so as to communicate with the front cooling passage 7 .
  • the cooling air CA which flows into the front cooling passage 7 is jetted out through the film cooling holes 11 .
  • the stator vane body 3 has a vane surface 3 v in its flank portion. Multiple film cooling holes 13 are formed in the vane surface 3 v so as to communicate with the front cooling passage 7 or the rear cooling passage 9 .
  • the cooling air CA which flows into the front cooling passage 7 or the rear cooling passage 9 is jetted out through the film cooling holes 13 .
  • the stator vane body 3 further has a vane surface 3 p in its rear edge-side portion. Multiple film cooling holes 15 are formed in the vane surface 3 p so as to communicate with the rear cooling passage 9 .
  • the cooling air CA which flows into the rear cooling passage 9 is jetted out through the film cooling holes 15 .
  • the stator vane body 3 has a vane surface 3 b in its back portion.
  • Multiple film cooling holes (not shown) similar to the film cooling holes 11 , the film cooling holes 13 and the film cooling holes 15 may be formed in the vane surface 3 b .
  • the film cooling holes are formed communicating with the front cooling passage 7 or the rear cooling passage 9 in the same way that has been described. The cooling air which flows into the front cooling passage 7 or the rear cooling passage 9 is jetted out through the film cooling holes (not shown).
  • an arc-shaped inner band 17 is integrally formed on a base-side portion of the stator vane body 3 in a span direction SD.
  • an arc-shaped outer band 19 is integrally formed on a tip-side portion of the stator vane body 3 in the span direction SD.
  • a front insertion hole 21 is formed in the outer band 19 so as to be aligned with the front cooling passage 7 of the stator vane body 3 .
  • a rear insertion hole 23 is formed in the outer band 19 so as to be aligned with the rear cooling passage 9 of the stator vane body 3 .
  • the front insertion hole 21 is formed in the front half of the outer band 19
  • the rear insertion hole 23 is formed in the rear half of the outer band 19 .
  • the front insertion hole 21 communicates with the front cooling passage 7 of the stator vane body 3
  • the rear insertion hole 23 communicates with the rear cooling passage 9 of the stator vane body 3 .
  • a pipe-shaped front insert 25 is arranged in the front cooling passage 7 of the stator vane body 3 .
  • the upper portion of the front insert 25 is inserted into the front insertion hole 21 of the outer band 19 .
  • Multiple front impingement cooling holes 27 are formed in the outer peripheral surface of the front insert 25 .
  • the cooling air CA is jetted out to the inner wall surface of the front cooling passage 7 through the front impingement cooling holes 27 .
  • a pipe-shaped rear insert 29 is arranged in the rear cooling passage 9 of the stator vane body 3 .
  • the upper portion of the rear insert 29 is inserted into the rear insertion hole 23 of the outer band 19 .
  • Multiple rear impingement cooling holes 31 are formed in the outer peripheral surface of the rear insert 29 .
  • the cooling air CA is jetted out to the inner wall surface of the rear cooling passage 9 of the stator vane body 3 through the rear impingement cooling holes 31 .
  • an entrance-side opening and an exit-side opening of each film cooling hole 11 have a rectangular long-hole shape extending in the span direction SD, and having rounded corners.
  • the hole cross section of each of the film cooling holes 11 has a rectangular long-hole shape extending in a direction PD parallel to the cross section along the span direction SD, and having rounded corners. It should be noted that the cross section shown in FIG. 4B is perpendicular to a hole-center line 11 c of the film cooling hole 11 .
  • the aspect ratio of the hole cross section of each film cooling hole 11 (or the ratio of the long side to the short side of the long hole viewed on the assumption that there are no rounded corners) is set in a range of 1.1 to 3.0, and preferably in a range of 1.5 to 2.5.
  • the reason for this is that: if the aspect ratio is less than 1.1, it is difficult to fully diffuse the cooling air CA, which is jetted out through each film cooling hole 11 , on the vane surface 3 a in the span direction SD; and if the aspect ratio is greater than 3.0, it is difficult to process each film cooling hole 11 (for example, by electrical discharge machining).
  • the hole-center line 11 c of each film cooling hole 11 tilts from a thickness direction TD of the stator vane body 3 on the cross section along the span direction SD (the cross section parallel to the span direction SD).
  • a tilt angle ⁇ c of each hole-center line 11 c from the thickness direction TD is set in a range of 20° to 60° or preferably in a range of 30° to 50°.
  • the reason for this is that: if the tilt angle ⁇ c is less than 20°, it is difficult to process the film cooling hole 11 , and stress concentration increases around the film cooling hole 11 ; and if the tilt angle ⁇ c is greater than 60°, the cooling air CA clearly comes up away (separates) from the stator vane body 3 , and the cooling performance accordingly deteriorates.
  • An exit-side and obtuse angle-side portion lie on the hole wall surface of each film cooling hole 11 tilts further to the span direction SD (or the vane surface 3 c ) than the hole-center line 11 c of the film cooling hole 11 on the cross section along the span direction SD.
  • a tilt angle ⁇ e of the predetermined exit-side portion 11 e of the hole wall surface of each film cooling hole 11 to the span direction SD (or the vane surface 3 a ) is in a range of 5° to 20° or preferably in a range of 5° to 10°.
  • the reason for this is that: if the tilt angle ⁇ e is less than 5°, it is not possible to fully diffuse the cooling air CA in the span direction SD; and if the tilt angle ⁇ e is greater than 20°, the stream is more likely to separate inside the hole wall surface of the film cooling hole 11 , and the cooling performance accordingly deteriorates.
  • the cooling air CA flowing into the front insert 25 is jetted out to the inner wall surface of the front cooling passage 7 of the stator vane body 3 through the multiple front impingement cooling holes 27
  • the cooling air CA flowing into the rear insert 29 is jetted out to the inner wall surface of the rear cooling passage 9 of the stator vane body 3 through the multiple rear impingement cooling holes 31 .
  • the impingement cooling inner cooling
  • the impingement cooling can be performed on the turbine stator vane 1 (the stator vane body 3 ).
  • the cooling air CA contributing to the impingement cooling of the turbine stator vane 1 in other words, the cooling air CA flowing into the front cooling passage 7 and the rear cooling passage 9 of the stator vane body 3 is jetted out through the multiple film cooling holes 11 , the multiple film cooling holes 13 and the multiple film cooling holes 15 .
  • the cooling air CA forms a film cooling layer (not shown) covering the front edge-side vane surface 3 a , the flank-side vane surface 3 v , and the like of the stator vane body 3 of the turbine stator vane 1 , and can perform the film cooling on the turbine stator vane 1 .
  • FIG. 5A is a cross-sectional view of a front edge-side vane surface 41 a of a turbine vane 41 of the embodiment of the present invention which was simulated in the above-mentioned analysis, the cross-sectional view taken along a span direction S 1 .
  • FIG. 5B is a cross-sectional view taken along a VB-VB line in FIG. 5A . This cross section is perpendicular to a hole-center line 43 c of a film cooling hole 43 .
  • the film cooling hole 43 is formed in the front edge-side vane surface 41 a of the turbine vane 41 (see FIG. 8 ).
  • the hole-center line 43 c of the film cooling hole 43 tilts from a thickness direction T 1 of the turbine vane 41 at the above-mentioned tilt angle ⁇ c on the cross section along the span direction S 1 .
  • an exit-side and obtuse angle-side portion (predetermined exit-side portion) of the hole wall surface of the film cooling hole 43 tilts to the span direction S 1 at the above-mentioned tilt angle ⁇ e, and tilts further to the span direction S 1 than the hole center line 43 c .
  • the film cooling hole 43 has a substantially rectangular cross section extending in a direction P 1 in parallel to the cross section along the span direction S 1 .
  • the film cooling hole 43 is a substantially rectangular long hole extending in the direction P 1 .
  • FIG. 6A is a cross-sectional view of a front edge-side vane surface 51 a of a turbine vane 51 (see FIG. 8 ) as Comparative Example 1 which was simulated in the above-mentioned analysis, the cross-sectional view taken along a span direction S 2 .
  • FIG. 6B is a cross-sectional view of the front edge-side vane surface 51 a taken along a VIB-VIB line in FIG. 6A . It should be noted that this cross section is perpendicular to a hole-center line 53 c of a film cooling hole 53 .
  • the film cooling hole 53 is formed in the front edge-side vane surface 51 a of the turbine vane 51 .
  • the hole-center line 53 a of the film cooling hole 53 tilts from a thickness direction T 2 of the turbine vane 51 at the predetermined angle ⁇ c on the cross section along the span direction S 2 .
  • the film cooling hole 53 has a circular cross section on a plane perpendicular to the hole-center line 53 c . It should be noted that the area of this circle is equal to the cross-sectional area of the film cooling hole 43 shown in FIG. 5B .
  • FIG. 7A is a cross-sectional view of a front edge-side vane surface 61 a of a turbine vane 61 (see FIG. 8 ) as Comparative Example 2 which was simulated in the above-mentioned analysis, the cross-sectional view taken along a span direction S 3 .
  • FIG. 7B is a cross-sectional view of the front edge-side vane surface 61 a taken along a VIIB-VIIB line in FIG. 7A . It should be noted that this cross section is perpendicular to a hole-center line 63 c of a film cooling hole 63 .
  • the film cooling hole 63 is formed in the front edge-side vane surface 61 a of the turbine vane 61 .
  • the hole-center line 63 a of the film cooling hole 63 tilts from a thickness direction T 3 of the turbine vane 61 at the predetermined angle ⁇ c on the cross section along the span direction S 3 .
  • the film cooling hole 63 has an elliptical cross section having a major axis along a direction P 3 on a plane perpendicular to the hole-center line 63 c .
  • the direction P 3 is in parallel to a cross section taken along the span direction S 3 . It should be noted that the area of this ellipse is equal to the cross-sectional area of the film cooling hole 43 shown in FIG. 5B .
  • this analysis was carried out on the assumption that a position of a stagnation point SP was 0° when viewed from the center OR of the curvature of the vane surface 41 a of the turbine vane 41 , and the exit-side opening portion of the film cooling hole 43 was formed in a position 71 which was displaced downstream of the stagnation point SP at 55° in the direction of a main flow F of a combustion gas. Furthermore, in this analysis, the front edge-side vane surfaces 41 a , 51 a , 61 a of the respective turbine vanes 41 , 51 , 61 were formed into the same shape.
  • FIGS. 9A to 9C , FIGS. 10A to 10C and FIGS. 11A to 11C show results of the analysis using the three-dimensional steady-state viscosity CFD.
  • FIGS. 9A to 9C show temperature distributions around the front edge-side vane surface of the turbine vane 41 of the embodiment.
  • FIGS. 10A to 10C show temperature distributions around the front edge-side vane surface of the turbine vane 51 as Comparative Example 1.
  • FIGS. 11A to 11C show temperature distributions around the front edge-side vane surface of the turbine vane 61 as Comparative Example 2.
  • FIG. 9A , FIG. 10A and FIG. 11A show the temperature distributions at the position 71 (see FIG.
  • FIG. 9B , FIG. 10B and FIG. 11B show the temperature distributions at a position 73 (see FIG. 8 ) which was displaced downstream of the stagnation point SP at 65°.
  • FIG. 9C , FIG. 10C and FIG. 11C show the temperature distributions at a position 75 (see FIG. 8 ) which was displaced downstream of the stagnation point SP at 75°.
  • the Reynolds number and the blow ratio were approximately 9 ⁇ 10 4 and 2.0, respectively, around the front edge-side vane surface.
  • an area H represents a high-temperature area corresponding to the combustion gas.
  • an area M represents an area where the mixture of the combustion gas and the cooling gas jetted out through the film cooling hole 43 ( 53 , 63 ) does not progress so much. In other words, the heat exchange does not progress fully between the combustion gas and the cooling gas in the area M. For this reason, the temperature of the area M is lower than that of the area H.
  • an area L represents an area where the mixture of the cooling gas and the combustion gas progresses even less than the area M. For this reason, the temperature of the area L is lower than that of the area M.
  • FIG. 12 shows a result of an analysis on average film efficiencies of the cooling gas around the front edge-side vane surface along the span direction which was obtained from another three-dimensional steady-state viscosity CFD analysis.
  • the film efficiency is defined as (Tg ⁇ Tf)/(Tg ⁇ Tc), where: Tg denotes the temperature of the combustion gas; Tf denotes the temperature of the film cooling layer; and Tc denotes the temperature of the cooling air CA.
  • the conditions for this analysis are the same as those for the foregoing analysis. That is to say, it was assumed that the Reynolds number and the blow ratio were approximately 9 ⁇ 10 4 and 2.0, respectively, around the front edge-side vane surface.
  • a solid line indicates a change in average film efficiency ⁇ 1 of the turbine vane 41 (see FIG. 5A and FIG. 5B ) of the embodiment which was used in the analysis.
  • a dotted line indicates a change in average film efficiency ⁇ 2 of the turbine vane 51 (see FIG. 6A and FIG. 6B ) of Comparative Example 1.
  • a chain dashed line indicates a change in average film efficiency ⁇ 3 of the turbine vane 61 (see FIG. 7A and FIG. 7B ) of Comparative Example 2.
  • the horizontal axis of the graph shown in FIG. 12 represents the distance of the front edge-side vane surface from the stagnation point SP.
  • the distance is converted into a dimensionless distance by dividing the distance by the equivalent diameter of the hole cross section of the film cooling hole (the diameter of a circular hole having the same cross-sectional area). It can be learned from this graph that: the average film efficiency ⁇ 1 of the turbine vane 41 of the embodiment is sufficiently higher than the average film efficiencies ⁇ 2 , ⁇ 3 of the turbine vanes 51 , 61 of Comparative Examples 1 and 2; and this tendency can be obtained regardless of how long the distance is.
  • the hole cross section of each film cooling hole 11 is shaped like a long hole extending in the direction PD parallel to the cross section along the span direction SD.
  • the hole-center line 11 c of each film cooling hole 11 tilts from the thickness direction TD on the cross section along the span direction SD.
  • the predetermined exit-side portion 11 e of the hole wall surface of each film cooling hole 11 tilts further from the thickness direction TD than the hole-center line 11 c on the cross section along the span direction SD.
  • each film cooling hole 11 is capable of more fully diffusing the cooling air CA, which is jetted out through the film cooling hole 11 , on the vane surface 3 a of the stator vane body 3 in the span direction SD while inhibiting the cooling air CA from coming up (separating) away from the vane surface 3 a than the case where the cross-sectional of the film cooling hole 11 would be shaped like a circle or an ellipse.
  • the embodiment makes it possible to increase the film efficiency around the vane surface 3 a of the stator vane body 3 of the turbine stator vane 1 , and accordingly to increase the cooling performance of the turbine stator vane 1 to a higher level.
  • the present invention is not limited to what has been described for the embodiment.
  • the present invention can be carried out in other various modes including, for example, a mode where the technical idea applied to the turbine stator vane 1 is applied to a turbine rotor vane. What is more, the scope of the right included in the present invention is not limited to these embodiments.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/046,191 2011-04-07 2013-10-04 Turbine vane Abandoned US20140037429A1 (en)

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JP2011085580A JP2012219702A (ja) 2011-04-07 2011-04-07 タービン翼
JP2011-085580 2011-04-07
PCT/JP2012/059459 WO2012137898A1 (ja) 2011-04-07 2012-04-06 タービン翼

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US20170101932A1 (en) * 2014-05-29 2017-04-13 General Electric Company Engine components with impingement cooling features
US20180051570A1 (en) * 2016-08-22 2018-02-22 Doosan Heavy Industries & Construction Co., Ltd. Gas turbine blade
US20190063243A1 (en) * 2016-02-24 2019-02-28 Safran Aircraft Engines Rectifier for aircraft turbomachine compressor, comprising air extraction openings having a stretched form in the peripheral direction
US10273812B2 (en) 2015-12-18 2019-04-30 Pratt & Whitney Canada Corp. Turbine rotor coolant supply system
US20190211686A1 (en) * 2018-01-05 2019-07-11 United Technologies Corporation Gas turbine engine airfoil with cooling path
US20190211687A1 (en) * 2018-01-05 2019-07-11 United Technologies Corporation Airfoil with rib communication
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WO2012137898A1 (ja) 2012-10-11
JP2012219702A (ja) 2012-11-12
EP2696030B1 (de) 2019-07-31
EP2696030A4 (de) 2014-10-29
CA2831382C (en) 2015-11-24
CA2831382A1 (en) 2012-10-11

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