US2956773A - Cooled hollow turbine blades - Google Patents

Cooled hollow turbine blades Download PDF

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Publication number
US2956773A
US2956773A US659375A US65937557A US2956773A US 2956773 A US2956773 A US 2956773A US 659375 A US659375 A US 659375A US 65937557 A US65937557 A US 65937557A US 2956773 A US2956773 A US 2956773A
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Prior art keywords
blade
cooling medium
partition
leading edge
passage
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US659375A
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Michael J French
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Napier Turbochargers Ltd
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D Napier and Son Ltd
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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
    • 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

Definitions

  • This invention relates to cooled hollow turbine blades. It is known to cool hollow turbine blades by passing a cooling medium, such as relatively cool air, through their hollow interiors, this air taking up heat from the walls of the hollow blades and thereby cooling them. In cooled blades which have hitherto been proposed it has been the usual practice to cause the cooling medium to flow predominantly in the longitudinal direction along the internal surfaces of the blades.
  • a cooling medium such as relatively cool air
  • the cooling medium is caused to flow across the internal surface of the blade adjacent to the leading edge in a direction approximately transverse to the length of the blade.
  • the leading edge of the blade is usually the region that needs the most cooling, and it has been found that by adopting an approximately transverse flow of cooling medium in accordance with the present invention the cooling of the leading edge is considerably improved. While the invention is not dependent upon any particular theory, it is thought that this improvement is due largely to the cooling medium sweeping quickly round the sharp curvature of the internal surface of the blade adjacent to the leading edge and removing the boundary layer of cooling medium from this surface more effectively than is the case when the flow is predominantly in the longitudinal direction.
  • the blade includes an internal member defining a longitudinal passage within the blade through which a cooling medium is passed from the blade root, and this passage has a slit or apertures disposed adjacent the internal surface of the blade in the vicinity of the leading edge and through which the cooling medium emerges from the passage in a direction approximately transverse to the length of the blade.
  • fresh cooling medium can be supplied along practically the whole length of the leading edge, providing substantially uniform cool-ing along this edge.
  • the cooling medium would get progressively hotter as it advanced along the leading edge, and in consequence would provide progressively less effective cooling.
  • the longitudinal passage can be formed in various ways. For instance, it may be bounded on one side by the internal member which in this case will be a partition dividing the hollow interior of the blade chordwise, and on the other side by one flank of the blade, preferably the convex flank.
  • the internal member may have a cross-section such that at least one of its sides is spaced by a small clearance from the adjacent flank of the blade so as to define at least one restricted passage for the flow of cooling medium, thereby ensuring that the cooling medium passes close to the internal surface of the blade on the flank or flanks as well as adjacent to the leading edge.
  • the internal member may be spaced from the flanks of the blade by forming raised pimples or weals thereon which touch the flanks of the blade.
  • the blade has outlet apertures for the cooling medium along its trailing edge, so that when the cooling medium has left the longitudinal chamber in the blade it flows substantially transversely, not only adjacent the leading edge but also across one or both flanks of the blade.
  • the blade instead of cooling medium emerging through the trailing edge of the blade, the blade may be provided with an internal return passage for the cooling medium leading back to the blade root or platform which will be provided with an outlet for the cooling medium.
  • Figure 1 is a perspective view of the blade, with the tip broken aawy to show the internal construction
  • Figure 2 is a cross-section taken on the line II-II in Figure 1;
  • Figure 3 is a fragmentary sectional view taken on the line III-III in Figure 2;
  • Figure 4 is a view, similar to Figure 3, of a modification.
  • Figure 5 is a cross-section, similar to Figure 2, of a further modification.
  • the hollow turbine blade shown in Figures 1 to- 3 comprises an aerofoil portion 10, a root portion 11 by which the blade is attached to the turbine rotor, and a platform portion 12.
  • the aerofoil portion 10 has a leading edge 13, a trailing edge 14, a convex flank 15 and a concave flank 16.
  • Secured within the hollow interior of the aerofoil portion 10 is a sheet metal partition 17, the partition having a row of apertures 18 down its forward edge and the tongues of metal 19 between these apertures being bent towards the convex flank 15 to touch the interior surface of this flank adjacent the leading edge 13.
  • the rear edge 20 of the partition touches the interior surface of the convex flank 15 towards the rear of the latter.
  • the platform portion 12 is provided with a cooling medium inlet aperture 21, and an internal baffle (not visible) which is conveniently an extension of the lower end of the partition 17 directs the flow of cooling medium as indicated by the dotted line provided with arrows into a longitudinal passage 22 defined on one side by the partition 17 and on the other side by the interior surface of the convex flank 15.
  • the other side of the partition 17 and the interior surface of the concave flank 16 together define another pas sage 23 which communicates with the passage 22 through the apertures 18.
  • the trailing edge 14 is provided with a series of apertures 24 communicating with the passage 23.
  • the partition 30 is provided, adjacent its forward edge, with raised pimples 31 which contact the convex blade flank 15 and provide apertures 32 corresponding to the apertures 18 of Figures 1-3.
  • the partition 17 is provided with a backing piece 33 which reduces the cross-sectional area between the partition 17 and flank 16 to define a restricted passage 23 and so provides a greater coolant flow velocity in this passage, thereby improving the cooling of the concave flank 16.
  • a backing piece 33 which reduces the cross-sectional area between the partition 17 and flank 16 to define a restricted passage 23 and so provides a greater coolant flow velocity in this passage, thereby improving the cooling of the concave flank 16.
  • all parts in this modification are identical to the corresponding parts of the preferred embodiment shown in Figures 1 and 2 and are, therefore, designated by similar reference characters.
  • a hollow turbine blade having a root portion, inlet means for a cooling medium in said root portion, a hollow aerofoil blade portion having a leading edge, a trailing edge provided with cooling medium outlet apertures, a convex flank and a concave flank, a partition extending longitudinally through the interior of said hollow aerofoil blade portion said front edge of said partition being adjacent said leading edge of said hollow aerofoil blade portion and said rear edge of said partition being in contact throughout its length with one of said flanks forwardly of said trailing edge, said partition defining between itself and said one flank a longitudinal passage extending lengthwise of said aerofoil portion and communicating with said cooling medium inlet means, said partition defining between itself and the other of said flanks a second passage communicating with said cooling medium outlet apertures, said partition being imperforate and separating said passages except adjacent said leading edge, and being formed adjacent said leading edge with at least one aperture for directing all of the said cooling medium from said first passage into said second passage at a location adjacent said leading edge and in a direction substantially

Description

Oct. 18, 1960 M. J. FRENCH COOLED HOLLOW TURBINE BLADES Filed May 15, .1957
INVENTOR MlcunsL J. FRENCH P M, MW
United States Patent O COOLED HOLLOW TURBINE BLADES Michael J. French, New Malden, England, assignor to D. Napier & Son Limited, Acton Vale, London, England, a company of Great Britain Filed May 15, 1957, Ser. No. 659,375
Claims priority, application Great Britain May 15, 1956 Claims. (Cl. 253-39.15)'
' This invention relates to cooled hollow turbine blades. It is known to cool hollow turbine blades by passing a cooling medium, such as relatively cool air, through their hollow interiors, this air taking up heat from the walls of the hollow blades and thereby cooling them. In cooled blades which have hitherto been proposed it has been the usual practice to cause the cooling medium to flow predominantly in the longitudinal direction along the internal surfaces of the blades.
According to the present invention, in a hollow turbine blades with provision for passing a cooling medium through it the cooling medium is caused to flow across the internal surface of the blade adjacent to the leading edge in a direction approximately transverse to the length of the blade.
The leading edge of the blade is usually the region that needs the most cooling, and it has been found that by adopting an approximately transverse flow of cooling medium in accordance with the present invention the cooling of the leading edge is considerably improved. While the invention is not dependent upon any particular theory, it is thought that this improvement is due largely to the cooling medium sweeping quickly round the sharp curvature of the internal surface of the blade adjacent to the leading edge and removing the boundary layer of cooling medium from this surface more effectively than is the case when the flow is predominantly in the longitudinal direction.
In one form of the invention the blade includes an internal member defining a longitudinal passage within the blade through which a cooling medium is passed from the blade root, and this passage has a slit or apertures disposed adjacent the internal surface of the blade in the vicinity of the leading edge and through which the cooling medium emerges from the passage in a direction approximately transverse to the length of the blade. Thus fresh cooling medium can be supplied along practically the whole length of the leading edge, providing substantially uniform cool-ing along this edge. On the other hand, with predominantly longitudinal flow the cooling medium would get progressively hotter as it advanced along the leading edge, and in consequence would provide progressively less effective cooling.
The longitudinal passage can be formed in various ways. For instance, it may be bounded on one side by the internal member which in this case will be a partition dividing the hollow interior of the blade chordwise, and on the other side by one flank of the blade, preferably the convex flank. The internal member may have a cross-section such that at least one of its sides is spaced by a small clearance from the adjacent flank of the blade so as to define at least one restricted passage for the flow of cooling medium, thereby ensuring that the cooling medium passes close to the internal surface of the blade on the flank or flanks as well as adjacent to the leading edge.
The internal member may be spaced from the flanks of the blade by forming raised pimples or weals thereon which touch the flanks of the blade.
ICC.
Preferably the blade has outlet apertures for the cooling medium along its trailing edge, so that when the cooling medium has left the longitudinal chamber in the blade it flows substantially transversely, not only adjacent the leading edge but also across one or both flanks of the blade. Alternatively, instead of cooling medium emerging through the trailing edge of the blade, the blade may be provided with an internal return passage for the cooling medium leading back to the blade root or platform which will be provided with an outlet for the cooling medium.
The invention may be performed in various ways, and one particular form of hollow turbine blade, and some modifications thereof, all embodying the invention, will now be described by way of example with reference to the accompanying drawings, in which:
Figure 1 is a perspective view of the blade, with the tip broken aawy to show the internal construction;
Figure 2 is a cross-section taken on the line II-II in Figure 1;
Figure 3 is a fragmentary sectional view taken on the line III-III in Figure 2;
Figure 4 is a view, similar to Figure 3, of a modification; and
Figure 5 is a cross-section, similar to Figure 2, of a further modification.
The hollow turbine blade shown in Figures 1 to- 3 comprises an aerofoil portion 10, a root portion 11 by which the blade is attached to the turbine rotor, and a platform portion 12. The aerofoil portion 10 has a leading edge 13, a trailing edge 14, a convex flank 15 and a concave flank 16. Secured within the hollow interior of the aerofoil portion 10 is a sheet metal partition 17, the partition having a row of apertures 18 down its forward edge and the tongues of metal 19 between these apertures being bent towards the convex flank 15 to touch the interior surface of this flank adjacent the leading edge 13. The rear edge 20 of the partition touches the interior surface of the convex flank 15 towards the rear of the latter. The platform portion 12 is provided with a cooling medium inlet aperture 21, and an internal baffle (not visible) which is conveniently an extension of the lower end of the partition 17 directs the flow of cooling medium as indicated by the dotted line provided with arrows into a longitudinal passage 22 defined on one side by the partition 17 and on the other side by the interior surface of the convex flank 15. The other side of the partition 17 and the interior surface of the concave flank 16 together define another pas sage 23 which communicates with the passage 22 through the apertures 18. The trailing edge 14 is provided with a series of apertures 24 communicating with the passage 23.
As the cooling medium advances longitudinally along the passage 22 portions of it escape through the apertures 18. These portions flow transversely across the interior surface of the leading edge 13, then flow transversely through the passage 23 and escape through the apertures 24 in the trailing edge. Thus all parts of the aerofoil portion are cooled by the cooling medium, the cooling eifect being particularly great adjacent the leading edge 13 owing, it is believed, to the sweeping effect of the curved transverse flow of the coolant providing effective coolant boundary layer removal in this region.
In the modification shown in Figure 4 the partition 30 is provided, adjacent its forward edge, with raised pimples 31 which contact the convex blade flank 15 and provide apertures 32 corresponding to the apertures 18 of Figures 1-3.
In the modification shown in Figure 5 the partition 17 is provided with a backing piece 33 which reduces the cross-sectional area between the partition 17 and flank 16 to define a restricted passage 23 and so provides a greater coolant flow velocity in this passage, thereby improving the cooling of the concave flank 16. Except as to the addition of the back piece 33, all parts in this modification are identical to the corresponding parts of the preferred embodiment shown in Figures 1 and 2 and are, therefore, designated by similar reference characters.
What I claim as my invention and desire to secure by Letters Patent is:
1. A hollow turbine blade having a root portion, inlet means for a cooling medium in said root portion, a hollow aerofoil blade portion having a leading edge, a trailing edge provided with cooling medium outlet apertures, a convex flank and a concave flank, a partition extending longitudinally through the interior of said hollow aerofoil blade portion said front edge of said partition being adjacent said leading edge of said hollow aerofoil blade portion and said rear edge of said partition being in contact throughout its length with one of said flanks forwardly of said trailing edge, said partition defining between itself and said one flank a longitudinal passage extending lengthwise of said aerofoil portion and communicating with said cooling medium inlet means, said partition defining between itself and the other of said flanks a second passage communicating with said cooling medium outlet apertures, said partition being imperforate and separating said passages except adjacent said leading edge, and being formed adjacent said leading edge with at least one aperture for directing all of the said cooling medium from said first passage into said second passage at a location adjacent said leading edge and in a direction substantially transverse to the length of said blade.
2. A hollow turbine blade according to claim 1 in which said partition has a cross-section such that said front edge thereof provides at least one aperture adjacent said leading edge of said aerofoil blade portion defining at least one restricted passage for the flow of cooling medium from said longitudinal passage.
3. A hollow turbine blade according to claim 1 in which said partition has raised portions which contact one of said flanks and locate said partition in said aerofoil portion.
4. A hollow turbine blade according to claim 1 in which said partition consists of a single thickness of sheet metal.
5. A hollow turbine blade according to claim 1 in which said partition consists of a hollow body made of sheet metal.
References Cited in the file of this patent UNITED STATES PATENTS 2,514,105 Thomas July 4, 1950 2,556,736 Palmaticr June 12, 1951 FOREIGN PATENTS 163,905 Australia July 5, 1955 920,641 Germany Nov. 25, 1954
US659375A 1956-05-15 1957-05-15 Cooled hollow turbine blades Expired - Lifetime US2956773A (en)

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GB (1) GB834811A (en)

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3220697A (en) * 1963-08-30 1965-11-30 Gen Electric Hollow turbine or compressor vane
US3420502A (en) * 1962-09-04 1969-01-07 Gen Electric Fluid-cooled airfoil
US3529902A (en) * 1968-05-22 1970-09-22 Gen Motors Corp Turbine vane
US3844678A (en) * 1967-11-17 1974-10-29 Gen Electric Cooled high strength turbine bucket
US3963368A (en) * 1967-12-19 1976-06-15 General Motors Corporation Turbine cooling
US4221539A (en) * 1977-04-20 1980-09-09 The Garrett Corporation Laminated airfoil and method for turbomachinery
US4416585A (en) * 1980-01-17 1983-11-22 Pratt & Whitney Aircraft Of Canada Limited Blade cooling for gas turbine engine
US5002460A (en) * 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade
US5165852A (en) * 1990-12-18 1992-11-24 General Electric Company Rotation enhanced rotor blade cooling using a double row of coolant passageways
EP0661414A1 (en) * 1993-12-28 1995-07-05 Kabushiki Kaisha Toshiba A cooled turbine blade for a gas turbine
US5704763A (en) * 1990-08-01 1998-01-06 General Electric Company Shear jet cooling passages for internally cooled machine elements
EP1207269A1 (en) * 2000-11-16 2002-05-22 Siemens Aktiengesellschaft Gas turbine vane
US20050169752A1 (en) * 2003-10-24 2005-08-04 Ching-Pang Lee Converging pin cooled airfoil
US20090148269A1 (en) * 2007-12-06 2009-06-11 United Technologies Corp. Gas Turbine Engines and Related Systems Involving Air-Cooled Vanes
US20120014808A1 (en) * 2010-07-14 2012-01-19 Ching-Pang Lee Near-wall serpentine cooled turbine airfoil
US20160072141A1 (en) * 2013-04-24 2016-03-10 Intelligent Energy Limited A water separator
US20160326885A1 (en) * 2015-05-08 2016-11-10 United Technologies Corporation Turbine engine component including an axially aligned skin core passage interrupted by a pedestal
US10344619B2 (en) * 2016-07-08 2019-07-09 United Technologies Corporation Cooling system for a gaspath component of a gas powered turbine
WO2019245546A1 (en) * 2018-06-20 2019-12-26 Siemens Energy, Inc. Cooled turbine blade assembly, corresponding methods for cooling and manufacturing
CN112459849A (en) * 2020-10-27 2021-03-09 哈尔滨广瀚燃气轮机有限公司 Cooling structure for turbine blade of gas turbine

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB895077A (en) * 1959-12-09 1962-05-02 Rolls Royce Blades for fluid flow machines such as axial flow turbines
US3301526A (en) * 1964-12-22 1967-01-31 United Aircraft Corp Stacked-wafer turbine vane or blade

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means
US2556736A (en) * 1945-06-22 1951-06-12 Curtiss Wright Corp Deicing system for aircraft
DE920641C (en) * 1943-07-15 1954-11-25 Maschf Augsburg Nuernberg Ag Cooled hollow blade, especially for gas turbines

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE710289C (en) * 1938-02-08 1941-09-09 Bbc Brown Boveri & Cie Blade with a device for the formation of a boundary layer protecting against high temperatures and a method for producing this blade
DE853534C (en) * 1943-02-27 1952-10-27 Maschf Augsburg Nuernberg Ag Air-cooled gas turbine blade
DE852786C (en) * 1943-11-10 1952-10-20 Versuchsanstalt Fuer Luftfahrt Time-graded cooling air throughput through the blades of gas or exhaust gas turbines
GB680581A (en) * 1949-05-09 1952-10-08 Hermann Oestrich Means for cooling the blades of gas turbine engines
GB685769A (en) * 1949-11-22 1953-01-14 Rolls Royce Improvements relating to compressor and turbine blading

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE920641C (en) * 1943-07-15 1954-11-25 Maschf Augsburg Nuernberg Ag Cooled hollow blade, especially for gas turbines
US2556736A (en) * 1945-06-22 1951-06-12 Curtiss Wright Corp Deicing system for aircraft
US2514105A (en) * 1945-12-07 1950-07-04 Thomas Wilfred Airfoil conditioning means

Cited By (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3420502A (en) * 1962-09-04 1969-01-07 Gen Electric Fluid-cooled airfoil
US3220697A (en) * 1963-08-30 1965-11-30 Gen Electric Hollow turbine or compressor vane
US3844678A (en) * 1967-11-17 1974-10-29 Gen Electric Cooled high strength turbine bucket
US3963368A (en) * 1967-12-19 1976-06-15 General Motors Corporation Turbine cooling
US3529902A (en) * 1968-05-22 1970-09-22 Gen Motors Corp Turbine vane
US4221539A (en) * 1977-04-20 1980-09-09 The Garrett Corporation Laminated airfoil and method for turbomachinery
US4416585A (en) * 1980-01-17 1983-11-22 Pratt & Whitney Aircraft Of Canada Limited Blade cooling for gas turbine engine
US5002460A (en) * 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade
US5704763A (en) * 1990-08-01 1998-01-06 General Electric Company Shear jet cooling passages for internally cooled machine elements
US5165852A (en) * 1990-12-18 1992-11-24 General Electric Company Rotation enhanced rotor blade cooling using a double row of coolant passageways
EP0661414A1 (en) * 1993-12-28 1995-07-05 Kabushiki Kaisha Toshiba A cooled turbine blade for a gas turbine
US5538394A (en) * 1993-12-28 1996-07-23 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
EP1207269A1 (en) * 2000-11-16 2002-05-22 Siemens Aktiengesellschaft Gas turbine vane
US6572329B2 (en) 2000-11-16 2003-06-03 Siemens Aktiengesellschaft Gas turbine
US20050169752A1 (en) * 2003-10-24 2005-08-04 Ching-Pang Lee Converging pin cooled airfoil
US6981840B2 (en) * 2003-10-24 2006-01-03 General Electric Company Converging pin cooled airfoil
US10156143B2 (en) * 2007-12-06 2018-12-18 United Technologies Corporation Gas turbine engines and related systems involving air-cooled vanes
US20090148269A1 (en) * 2007-12-06 2009-06-11 United Technologies Corp. Gas Turbine Engines and Related Systems Involving Air-Cooled Vanes
US20120014808A1 (en) * 2010-07-14 2012-01-19 Ching-Pang Lee Near-wall serpentine cooled turbine airfoil
US8535006B2 (en) * 2010-07-14 2013-09-17 Siemens Energy, Inc. Near-wall serpentine cooled turbine airfoil
US20130302167A1 (en) * 2010-07-14 2013-11-14 Mikro Systems, Inc. Near-Wall Serpentine Cooled Turbine Airfoil
US8870537B2 (en) * 2010-07-14 2014-10-28 Mikro Systems, Inc. Near-wall serpentine cooled turbine airfoil
US20160072141A1 (en) * 2013-04-24 2016-03-10 Intelligent Energy Limited A water separator
US20160326885A1 (en) * 2015-05-08 2016-11-10 United Technologies Corporation Turbine engine component including an axially aligned skin core passage interrupted by a pedestal
US10502066B2 (en) * 2015-05-08 2019-12-10 United Technologies Corporation Turbine engine component including an axially aligned skin core passage interrupted by a pedestal
US11143039B2 (en) 2015-05-08 2021-10-12 Raytheon Technologies Corporation Turbine engine component including an axially aligned skin core passage interrupted by a pedestal
US10344619B2 (en) * 2016-07-08 2019-07-09 United Technologies Corporation Cooling system for a gaspath component of a gas powered turbine
WO2019245546A1 (en) * 2018-06-20 2019-12-26 Siemens Energy, Inc. Cooled turbine blade assembly, corresponding methods for cooling and manufacturing
CN112459849A (en) * 2020-10-27 2021-03-09 哈尔滨广瀚燃气轮机有限公司 Cooling structure for turbine blade of gas turbine

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GB834811A (en) 1960-05-11
CH342413A (en) 1959-11-15
DE1056427B (en) 1959-04-30
BE557503A (en)
FR1175169A (en) 1959-03-20

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