US20100143132A1 - Turbine blade with reverse cooling air film hole direction - Google Patents
Turbine blade with reverse cooling air film hole direction Download PDFInfo
- Publication number
- US20100143132A1 US20100143132A1 US12/706,777 US70677710A US2010143132A1 US 20100143132 A1 US20100143132 A1 US 20100143132A1 US 70677710 A US70677710 A US 70677710A US 2010143132 A1 US2010143132 A1 US 2010143132A1
- Authority
- US
- United States
- Prior art keywords
- section
- airfoil
- turbine blade
- meter
- set forth
- 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
Links
- 238000001816 cooling Methods 0.000 title claims abstract description 47
- 238000002485 combustion reaction Methods 0.000 description 5
- 239000000446 fuel Substances 0.000 description 2
- 238000000034 method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
Images
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
Abstract
A gas turbine engine includes turbine blades having film cooling holes at an outer face of an airfoil wherein the film cooling holes are designed to be better filled with air. In a disclosed embodiment, the film cooling holes include a meter section extending along a direction having a main component extending from a blade tip to a blade root. In addition, a diffused section communicates with the meter section at a face of the airfoil. The diffused section is spaced toward the blade tip from the meter section. In this manner, centrifugal force ensures the diffused section is also filled with air.
Description
- This application is a continuation of U.S. patent application Ser. No. 11/651,226, which was filed Jan. 9, 2007.
- This application relates to a turbine blade, wherein the meter sections of film cooling holes extend at an angle and in a direction toward a blade root from the blade tip. In addition, a diffused section of a film cooling hole extends toward the blade tip from a meter section to receive air driven by centrifugal force.
- Gas turbine engines are known, and include a plurality of sections which are typically mounted in series. Typically a fan delivers air to a compressor. Air is compressed in the compressor and delivered downstream to be mixed with fuel and combusted in a combustor section. Products of combustion move downstream over turbine rotors. The turbine rotors include a plurality of removable blades which rotate with the rotors, and are driven by the products of combustion. The turbine rotors drive components within the gas turbine engine, including the fan and compressor.
- The turbine blades become quite hot from the products of combustion. Thus, it is known to pass cooling air through internal cooling passages within the turbine blades. In one known cooling technique, air is passed outwardly through holes on an outer face of an airfoil of the turbine blade, such that the cool air passes along the outer face. These film cooling holes are designed to maximize the coverage surface area on the blade, which receives the air and also to maximize the time cooling air is kept on a face of the blade.
- In the prior art, the film cooling holes have a meter section that typically extend at an angle to the outer face. The angle includes a major component in a direction extending from a blade root and toward a blade tip. In addition, a diffused section extends back from this meter section towards the blade root. This type of film cooling holes is known as shaped or flared holes. The purpose of the diffused section is to slow the speed of the cooling air down as it reaches the face of the blade, such that the air would be less likely to move away from the face, and more likely to move along the face.
- However, in the prior art, a centrifugal force applied as the blade rotates, moves the cooling air radially outwardly and toward the blade tip. Thus, the diffused section tends not to be filled with air. This centrifugal force and flow momentum drives the air into the radially outer portions of the holes spaced toward the tip, and leaves the diffused section less filled. Thus, the air exits the film cooling hole at a greater velocity, and does not stay on the face of the blade as long as would be desired.
- In a disclosed embodiment of this invention, the meter section of film cooling holes in a turbine blade extend with a major component in a direction from the blade tip toward the blade root. A diffused section is formed to enlarge a film cooling hole at the outer face of the blade. The diffused section extends toward the blade tip from the meter section.
- As the blade rotates, and cooling air exits the film cooling hole, centrifugal force forces some of the cooling air into the diffused section and the diffused section is relatively full compared to the prior art. Thus, the air exits the film cooling hole at a lower velocity than in the prior art, tends to stay on the face of the turbine blade longer, and cover a greater surface area.
- These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
-
FIG. 1 is a schematic view of a gas turbine engine. -
FIG. 2A is a view of a prior art turbine blade. -
FIG. 2B is an enlarged view of a portion of theFIG. 2A turbine blade. -
FIG. 2C is another view of theFIG. 2A blade. -
FIG. 3 is a view similar toFIG. 2C , but showing the inventive blade. - A
gas turbine engine 10 circumferentially disposed about an engine centerline, oraxial centerline axis 12 is shown inFIG. 1 . Theengine 10 includes afan 14, acompressor 16, acombustion section 18 and aturbine 11. As is well known in the art, air compressed in thecompressor 16 is mixed with fuel and burned in thecombustion section 18 and expanded inturbine 11. Theturbine 11 includesrotors 22 which rotate in response to the expansion, driving thecompressor 16 andfan 14. Theturbine 11 comprises alternating rows of rotary airfoils orblades 24 and static airfoils orvanes 26. In fact, this view is quite schematic, andblades 24 andvanes 26 are actually removable. It should be understood that this view is included simply to provide a basic understanding of the sections in a gas turbine engine, and not to limit the invention. This invention extends to all types of turbine engines for all applications. -
FIG. 2A shows a priorart turbine blade 24. As known, aplatform 32 and blade root form a base for anairfoil 34. Theairfoil 34 includes a plurality offilm cooling holes 36. As can be appreciated, theholes 36 are formed on the pressure side 198 of the turbine blade. The holes are in an array, with holes being spaced in several columns and rows extending between the root of the airfoil and the tip, and from the trailing edge 197 toward the leading edge 199. As an example, there are several columns 200, 201, and 202 spaced between the trailing edge 197 and the leading edge 199. In addition, there are holes 202 that are closer to the root than other holes 205 or 207. - As shown in
FIG. 2B , thefilm cooling holes 36 have ameter section 38, and a diffusedsection 40. - As shown in
FIG. 2C , themeter section 38 extends along a non-parallel angle relative to a radial axis, and with a component extending from the blade root to the blade tip. The air from aninternal cooling passage 42 passes through thismeter section 38 to an outer face of theairfoil 34. As can be seen inFIG. 2C , this diffused section extends from themeter section 38 and closer to the blade root than the blade tip. Now, as theturbine blade 24 rotates, centrifugal forces force air from themeter section 38 radially outwardly, and away from the diffusedsection 40. Thus, the diffusedsection 40 is not always filled. - As shown in
FIG. 3 , in aninventive turbine blade 50, ameter section 52 extends with a main component of its direction from the blade tip to the blade root. A diffusedsection 54 extends toward the blade tip from themeter section 52. As can be seen, the diffusedsection 54 may be at an angle having a lesser component in the direction from the tip towards the root. As can be appreciated fromFIG. 2 , theenlarged portions FIG. 3 , themeter sections 52 extend from coolingpassage 42 at an angle that is initially from the blade tip toward the blade root, and at a single angle to an outer face of the airfoil. Whileholes 52 are shown along a single column, it should be appreciated that these holes would be utilized in an array such as shown inFIG. 2A or 2B. - When centrifugal force acts on the air in the
meter section 52, the air is driven into the diffusedsection 54. Flow momentum will ensure that themeter section 52 is still full. Thus, the present invention ensures the cooling air is delivered to theouter face 51 across the entirety of the film cooling holes. As can be appreciated fromFIG. 2B , the diffusedsections internal cooling passage 42 can flow in any direction and does not necessarily have to flow from blade root to blade tip. - In fact, the meter section can extend in the reverse direction or any direction with the diffused section extending toward the tip. Flow momentum will still fill the meter section while centrifugal force will fill the diffused section.
- Although a preferred embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
Claims (13)
1. A turbine blade comprising:
a root, and an airfoil extending away from the root to a tip;
a plurality of film cooling holes on an outer face of the airfoil, said airfoil having at least one internal cooling passage for receiving air from a source, and delivering air to said film cooling holes; and
said film cooling holes receiving air from said cooling passage through meter sections extending with a component in a direction from the tip towards the root.
2. The turbine blade as set forth in claim 1 , wherein a diffused section of an outer end of said film cooling holes extending towards said tip from said meter section.
3. The turbine blade as set forth in claim 2 , wherein said diffused section is formed along an angle having a lesser component in the direction from said tip toward said root than said meter section.
4. The turbine blade as set forth in claim 2 , wherein said meter sections extend at a first angle, with an extension of said meter section extending through to an outer wall of said airfoil, and said diffused section extending at a different angle than said meter section.
5. The turbine blade as set forth in claim 1 , wherein said plurality of film holes being formed in an array, with there being film holes spaced at different locations in a direction between a trailing edge and a leading edge of the airfoil, and also at different locations between the root and the tip of the airfoil.
6. The turbine blade as set forth in claim 1 , wherein said meter section extends from said cooling passage initially at said direction.
7. The turbine blade as set forth in claim 10 , wherein said meter section extends along a single angle from said cooling passage to said outer wall of said outer face of the airfoil.
8. The turbine blade as set forth in claim 1 , wherein said array of cooling holes is formed on a pressure side of said airfoil.
9. A turbine blade comprising:
a root, and an airfoil extending away from the root toward a tip;
a plurality of film cooling holes on an outer face of said airfoil, said airfoil having at least one internal cooling passage for receiving air from a source, and delivering air to said film cooling holes; and
said film cooling holes receiving air from said internal cooling passage through meter sections, and an diffused section of an outer end of said film cooling holes communicates with said meter section, said diffused section extending towards said tip from said meter section.
10. The turbine blade as set forth in claim 9 , wherein said diffused section is formed along an angle having a lesser component in the direction from said tip toward said root than said meter section.
11. The turbine blade as set forth in claim 9 , wherein said meter section extends from said cooling passage initially at said direction.
12. The turbine blade as set forth in claim 11 , wherein said meter sections each extend along a single angle from said cooling passage to said outer wall of said outer face of the airfoil.
13. The turbine blade as set forth in claim 9 , wherein said array of cooling holes is formed on a pressure side of said airfoil.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/706,777 US20100143132A1 (en) | 2007-01-09 | 2010-02-17 | Turbine blade with reverse cooling air film hole direction |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US11/651,226 US7712316B2 (en) | 2007-01-09 | 2007-01-09 | Turbine blade with reverse cooling air film hole direction |
US12/706,777 US20100143132A1 (en) | 2007-01-09 | 2010-02-17 | Turbine blade with reverse cooling air film hole direction |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/651,226 Continuation US7712316B2 (en) | 2007-01-09 | 2007-01-09 | Turbine blade with reverse cooling air film hole direction |
Publications (1)
Publication Number | Publication Date |
---|---|
US20100143132A1 true US20100143132A1 (en) | 2010-06-10 |
Family
ID=39267819
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/651,226 Active 2027-10-13 US7712316B2 (en) | 2007-01-09 | 2007-01-09 | Turbine blade with reverse cooling air film hole direction |
US12/706,777 Abandoned US20100143132A1 (en) | 2007-01-09 | 2010-02-17 | Turbine blade with reverse cooling air film hole direction |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/651,226 Active 2027-10-13 US7712316B2 (en) | 2007-01-09 | 2007-01-09 | Turbine blade with reverse cooling air film hole direction |
Country Status (2)
Country | Link |
---|---|
US (2) | US7712316B2 (en) |
EP (1) | EP1947296B1 (en) |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8245519B1 (en) * | 2008-11-25 | 2012-08-21 | Florida Turbine Technologies, Inc. | Laser shaped film cooling hole |
CN104281751A (en) * | 2014-10-14 | 2015-01-14 | 北京航空航天大学 | Feature-based parametric build system and method of turbine cooling blade |
CN104392027A (en) * | 2014-11-10 | 2015-03-04 | 西北工业大学 | Parametric modeling method of turbine blade turbulence flow column |
CN104598684A (en) * | 2015-01-19 | 2015-05-06 | 西北工业大学 | Parametric modeling method for film hole |
Families Citing this family (24)
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---|---|---|---|---|
US8105033B2 (en) * | 2008-06-05 | 2012-01-31 | United Technologies Corporation | Particle resistant in-wall cooling passage inlet |
US8210814B2 (en) * | 2008-06-18 | 2012-07-03 | General Electric Company | Crossflow turbine airfoil |
US8079810B2 (en) * | 2008-09-16 | 2011-12-20 | Siemens Energy, Inc. | Turbine airfoil cooling system with divergent film cooling hole |
US9121290B2 (en) * | 2010-05-06 | 2015-09-01 | United Technologies Corporation | Turbine airfoil with body microcircuits terminating in platform |
US9157328B2 (en) | 2010-12-24 | 2015-10-13 | Rolls-Royce North American Technologies, Inc. | Cooled gas turbine engine component |
US8533949B2 (en) * | 2011-02-14 | 2013-09-17 | General Electric Company | Methods of manufacture for components with cooling channels |
US20130156602A1 (en) * | 2011-12-16 | 2013-06-20 | United Technologies Corporation | Film cooled turbine component |
US8763402B2 (en) | 2012-02-15 | 2014-07-01 | United Technologies Corporation | Multi-lobed cooling hole and method of manufacture |
EP2956633B1 (en) | 2013-02-15 | 2021-05-05 | Raytheon Technologies Corporation | Component for a gas turbine engine and corresponding method of forming a cooling hole |
US9371776B2 (en) | 2013-08-20 | 2016-06-21 | Darren Levine | Dual flow air injection intraturbine engine and method of operating same |
US9416662B2 (en) * | 2013-09-03 | 2016-08-16 | General Electric Company | Method and system for providing cooling for turbine components |
US10036259B2 (en) | 2014-11-03 | 2018-07-31 | United Technologies Corporation | Turbine blade having film cooling hole arrangement |
US10443434B2 (en) | 2014-12-08 | 2019-10-15 | United Technologies Corporation | Turbine airfoil platform segment with film cooling hole arrangement |
US10301966B2 (en) | 2014-12-08 | 2019-05-28 | United Technologies Corporation | Turbine airfoil platform segment with film cooling hole arrangement |
US10107140B2 (en) | 2014-12-08 | 2018-10-23 | United Technologies Corporation | Turbine airfoil segment having film cooling hole arrangement |
US10060268B2 (en) | 2014-12-17 | 2018-08-28 | United Technologies Corporation | Turbine blade having film cooling hole arrangement |
US20160298545A1 (en) * | 2015-04-13 | 2016-10-13 | General Electric Company | Turbine airfoil |
US20170234142A1 (en) * | 2016-02-17 | 2017-08-17 | General Electric Company | Rotor Blade Trailing Edge Cooling |
US20170298743A1 (en) * | 2016-04-14 | 2017-10-19 | General Electric Company | Component for a turbine engine with a film-hole |
US10731469B2 (en) | 2016-05-16 | 2020-08-04 | Raytheon Technologies Corporation | Method and apparatus to enhance laminar flow for gas turbine engine components |
CN111706409B (en) * | 2020-06-25 | 2022-11-01 | 中国民航大学 | Corrugated air film hole with branch hole |
US11898460B2 (en) | 2022-06-09 | 2024-02-13 | General Electric Company | Turbine engine with a blade |
US11927111B2 (en) | 2022-06-09 | 2024-03-12 | General Electric Company | Turbine engine with a blade |
CN117226614B (en) * | 2023-11-14 | 2024-01-12 | 中国航发沈阳黎明航空发动机有限责任公司 | Method for polishing air film holes of double-wall turbine blade of aero-engine |
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US4384823A (en) * | 1980-10-27 | 1983-05-24 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Curved film cooling admission tube |
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US20080152475A1 (en) * | 2006-12-21 | 2008-06-26 | Jack Raul Zausner | Method for preventing backflow and forming a cooling layer in an airfoil |
US7621718B1 (en) * | 2007-03-28 | 2009-11-24 | Florida Turbine Technologies, Inc. | Turbine vane with leading edge fillet region impingement cooling |
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US3819295A (en) * | 1972-09-21 | 1974-06-25 | Gen Electric | Cooling slot for airfoil blade |
GB2401915B (en) * | 2003-05-23 | 2006-06-14 | Rolls Royce Plc | Turbine blade |
-
2007
- 2007-01-09 US US11/651,226 patent/US7712316B2/en active Active
-
2008
- 2008-01-08 EP EP08250077.8A patent/EP1947296B1/en active Active
-
2010
- 2010-02-17 US US12/706,777 patent/US20100143132A1/en not_active Abandoned
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US4384823A (en) * | 1980-10-27 | 1983-05-24 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Curved film cooling admission tube |
US4653983A (en) * | 1985-12-23 | 1987-03-31 | United Technologies Corporation | Cross-flow film cooling passages |
US4992025A (en) * | 1988-10-12 | 1991-02-12 | Rolls-Royce Plc | Film cooled components |
US5340278A (en) * | 1992-11-24 | 1994-08-23 | United Technologies Corporation | Rotor blade with integral platform and a fillet cooling passage |
US5419681A (en) * | 1993-01-25 | 1995-05-30 | General Electric Company | Film cooled wall |
US5503529A (en) * | 1994-12-08 | 1996-04-02 | General Electric Company | Turbine blade having angled ejection slot |
US5498133A (en) * | 1995-06-06 | 1996-03-12 | General Electric Company | Pressure regulated film cooling |
US6092982A (en) * | 1996-05-28 | 2000-07-25 | Kabushiki Kaisha Toshiba | Cooling system for a main body used in a gas stream |
US6164913A (en) * | 1999-07-26 | 2000-12-26 | General Electric Company | Dust resistant airfoil cooling |
US6234755B1 (en) * | 1999-10-04 | 2001-05-22 | General Electric Company | Method for improving the cooling effectiveness of a gaseous coolant stream, and related articles of manufacture |
US20080152475A1 (en) * | 2006-12-21 | 2008-06-26 | Jack Raul Zausner | Method for preventing backflow and forming a cooling layer in an airfoil |
US7621718B1 (en) * | 2007-03-28 | 2009-11-24 | Florida Turbine Technologies, Inc. | Turbine vane with leading edge fillet region impingement cooling |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8245519B1 (en) * | 2008-11-25 | 2012-08-21 | Florida Turbine Technologies, Inc. | Laser shaped film cooling hole |
CN104281751A (en) * | 2014-10-14 | 2015-01-14 | 北京航空航天大学 | Feature-based parametric build system and method of turbine cooling blade |
CN104392027A (en) * | 2014-11-10 | 2015-03-04 | 西北工业大学 | Parametric modeling method of turbine blade turbulence flow column |
CN104598684A (en) * | 2015-01-19 | 2015-05-06 | 西北工业大学 | Parametric modeling method for film hole |
Also Published As
Publication number | Publication date |
---|---|
EP1947296A3 (en) | 2014-01-15 |
US20080163604A1 (en) | 2008-07-10 |
US7712316B2 (en) | 2010-05-11 |
EP1947296A2 (en) | 2008-07-23 |
EP1947296B1 (en) | 2015-02-25 |
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Legal Events
Date | Code | Title | Description |
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STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |