US20110236223A1 - Blade for a gas turbine - Google Patents
Blade for a gas turbine Download PDFInfo
- Publication number
- US20110236223A1 US20110236223A1 US13/075,234 US201113075234A US2011236223A1 US 20110236223 A1 US20110236223 A1 US 20110236223A1 US 201113075234 A US201113075234 A US 201113075234A US 2011236223 A1 US2011236223 A1 US 2011236223A1
- Authority
- US
- United States
- Prior art keywords
- platform
- blade
- airfoil
- transition
- cooling medium
- 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
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
- F01D5/187—Convection 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
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
-
- 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
- F05D2250/20—Three-dimensional
- F05D2250/24—Three-dimensional ellipsoidal
-
- 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 the field of gas turbines, and particularly relates to a blade for a gas turbine.
- the requirements for increasing the efficiency of gas turbines leads to the thickness at the trailing edges of the blade airfoils of the blades which are fitted in the gas turbines having to be continuously reduced. This results in a geometry of the blade as is exemplarily shown in cross section in FIG. 1 .
- the blade 10 of FIG. 1 extends in the manner of an airfoil profile transversely to its longitudinal direction between a rounded leading edge 15 and a comparatively sharply tapering trailing edge 16 .
- the blade 10 has a (concave) pressure side 13 and a (convex) suction side 14 with corresponding walls 13 ′ and 14 ′.
- a gaseous cooling medium or coolant is fed in the hollow inner space 17 and discharged into the environment inter alia through a cooling medium outlet which is formed at the trailing edge 16 .
- a particularly sharply tapering, slender trailing edge 16 in this case is achieved by the cooling medium outlet 18 being arranged entirely on the pressure side 13 of the blade 10 , and by the two walls 13 ′ and 14 ′ being constructed especially thin in the region of the trailing edge 16 .
- the transition of the blade airfoil 11 to this platform 12 in the region of the trailing edge 16 represents a typical factor which limits the service life of a gas turbine component because it is exposed to superposition of high thermal stress which is brought about by the thermo-mechanical mismatch between platform 12 and blade airfoil 11 , and to mechanical stress peaks which are brought about by loading of the blades as a result of the gas flow. Reducing the thickness of the trailing edge 16 causes an increase of the stress in this critical region so that when designing the blade measures have to be considered in order to achieve and to ensure a sufficiently long service life.
- An aspect of the present invention is to further develop a blade for a gas turbine so that despite a low thickness at the trailing edge of the blade airfoil a satisfactory service life is achieved.
- the present invention provides a blade for a gas turbine including an airfoil extending in a longitudinal direction and extending transversely to the longitudinal direction between a leading edge and a trailing edge.
- the airfoil has a pressure side, a suction side, and a slot-like cooling medium outlet extending along the trailing edge.
- the cooling medium outlet is configured to discharge cooling medium supplied from an inner space of the blade.
- a platform extends transversely to the longitudinal direction.
- An end of the airfoil merges into an underside of the platform and has a transition from the airfoil to the platform at the trailing edge of the airfoil that increases in thickness in a direction toward the underside of the platform.
- the cooling medium outlet extends into the platform so as to reduce an operating temperature in a region of the transition from the blade airfoil to the platform.
- FIG. 1 shows a simplified cross section through an gas turbine blade with a narrow trailing edge and a cooling medium outlet
- FIG. 2 shows a sharp transition between blade airfoil and platform in a blade such as that shown in FIG. 1 ;
- FIG. 3 shows a low-stress transition between blade airfoil and platform according to an exemplary embodiment of the invention.
- the present invention provides a transition from the blade airfoil to the platform at the trailing edge that has a transition thickness profile in which the thickness increases above average the closer it gets to the underside of the platform, and that the cooling medium outlet is extended right into the platform for reducing the temperature in the region of the transition from the blade airfoil to the platform.
- a transition thickness profile in which the thickness increases above average the closer it gets to the underside of the platform, and that the cooling medium outlet is extended right into the platform for reducing the temperature in the region of the transition from the blade airfoil to the platform.
- the transition thickness profile has an essentially exponential shape which resembles an inverted, very slender, truncated pyramid or an inverted virtual pyramid. As a result of this an especially “smooth” transition between trailing edge and platform is achieved.
- the transition from the blade airfoil to the platform has an approximately elliptical transition border profile which also reduces stresses.
- the cooling medium outlet is formed between a pressure-side wall of the blade airfoil and a suction-side wall of the blade airfoil, and in that the pressure-side wall has a curvilinear transition edge profile in the transition from the blade airfoil to the platform in such a way that the wall thickness of the pressure-side wall in the region of the transition from the blade airfoil to the platform is approximately equal to the wall thickness in the remaining region of the blade airfoil.
- a blade 20 for a gas turbine with a low-stress transition between blade airfoil 11 and platform 12 according to an exemplary embodiment of the invention is reproduced.
- the blade 20 of the exemplary embodiment comprises a blade airfoil 11 which extends in a longitudinal direction and in the manner of a wing extends transversely to the longitudinal direction between a leading edge 15 and a trailing edge 16 , and has a pressure side 13 and a suction side 14 .
- the blade airfoil 11 merges into a platform 12 which lies transversely to the longitudinal direction and projects laterally across the blade cross section.
- a slot-like cooling medium outlet 18 which extends along the trailing edge 16 is provided at the trailing edge 16 of the blade airfoil 11 , through which a cooling medium, for example cooling air, which is fed via the (hollow) inner space 17 of the blade 20 , is discharged.
- the trailing edge 16 with its thin walls 13 ′ and 14 ′ is very narrow in construction.
- the transition has a transition thickness profile 21 in which the thickness D increases above average the closer it gets to the underside 12 ′ of the platform 12 .
- the cooling medium outlet 18 is extended (extension 19 ) right into the platform 12 for reducing the local temperature in the region of the transition from the blade airfoil 11 to the platform 12 .
- the transition thickness profile 21 has an essentially exponential shape and as a result resembles an inverted rampant pyramid. It is especially favorable in this case with regard to the stress distribution if the transition from the blade airfoil 11 to the platform 12 has an approximately elliptical transition border profile 22 . While in the case of the conventional blade according to FIG. 2 the trailing edge 16 of the blade airfoil 11 terminates inside the platform 12 and does not extend as far as the boundary of the platform 12 , in the case of the exemplary embodiment of FIG. 3 the trailing edge 16 at the transition from the blade airfoil 11 to the platform 12 is extended up to the edge 12 ′′ of the platform 12 .
- the cooling medium outlet 18 is delimited by the pressure-side wall 13 ′ and the suction-side wall 14 ′ of the blade airfoil 11 .
- the pressure-side wall 13 ′ in this case has a curvilinear transition edge profile 23 in the transition from the blade airfoil 11 to the platform 12 in such a way that the wall thickness of the pressure-side wall 13 ′ in the region of the transition from the blade airfoil 11 to the platform 12 is approximately equal to the wall thickness in the remaining region of the blade airfoil 11 .
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CH01548/08A CH699601A1 (de) | 2008-09-30 | 2008-09-30 | Schaufel für eine gasturbine. |
CH01548/08 | 2008-09-30 | ||
PCT/EP2009/062213 WO2010037659A1 (de) | 2008-09-30 | 2009-09-21 | Schaufel für eine gasturbine |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/EP2009/062213 Continuation WO2010037659A1 (de) | 2008-09-30 | 2009-09-21 | Schaufel für eine gasturbine |
Publications (1)
Publication Number | Publication Date |
---|---|
US20110236223A1 true US20110236223A1 (en) | 2011-09-29 |
Family
ID=40130846
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US13/075,234 Abandoned US20110236223A1 (en) | 2008-09-30 | 2011-03-30 | Blade for a gas turbine |
Country Status (4)
Country | Link |
---|---|
US (1) | US20110236223A1 (de) |
CH (1) | CH699601A1 (de) |
DE (1) | DE112009002628T5 (de) |
WO (1) | WO2010037659A1 (de) |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2868867A1 (de) * | 2013-10-29 | 2015-05-06 | Siemens Aktiengesellschaft | Turbinenschaufel |
US9045987B2 (en) | 2012-06-15 | 2015-06-02 | United Technologies Corporation | Cooling for a turbine airfoil trailing edge |
WO2018031032A1 (en) * | 2016-08-12 | 2018-02-15 | Siemens Aktiengesellschaft | Blade for gas turbine engine |
Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6062817A (en) * | 1998-11-06 | 2000-05-16 | General Electric Company | Apparatus and methods for cooling slot step elimination |
US6190128B1 (en) * | 1997-06-12 | 2001-02-20 | Mitsubishi Heavy Industries, Ltd. | Cooled moving blade for gas turbine |
US6328531B1 (en) * | 1998-08-05 | 2001-12-11 | Societe Nationale d'Etude et de Construction de Moteurs d'Aviation “SNECMA” | Cooled turbine blade |
US20030138322A1 (en) * | 2002-01-23 | 2003-07-24 | Snecma Moteurs | Moving blade for a high pressure turbine, the blade having a trailing edge of improved thermal behavior |
US6851924B2 (en) * | 2002-09-27 | 2005-02-08 | Siemens Westinghouse Power Corporation | Crack-resistance vane segment member |
US20050249593A1 (en) * | 2004-01-14 | 2005-11-10 | Snecma Moteurs | Cooling air evacuation slots of turbine blades |
US7249933B2 (en) * | 2005-01-10 | 2007-07-31 | General Electric Company | Funnel fillet turbine stage |
US7371046B2 (en) * | 2005-06-06 | 2008-05-13 | General Electric Company | Turbine airfoil with variable and compound fillet |
US7581924B2 (en) * | 2006-07-27 | 2009-09-01 | Siemens Energy, Inc. | Turbine vanes with airfoil-proximate cooling seam |
-
2008
- 2008-09-30 CH CH01548/08A patent/CH699601A1/de not_active Application Discontinuation
-
2009
- 2009-09-21 DE DE112009002628T patent/DE112009002628T5/de active Pending
- 2009-09-21 WO PCT/EP2009/062213 patent/WO2010037659A1/de active Application Filing
-
2011
- 2011-03-30 US US13/075,234 patent/US20110236223A1/en not_active Abandoned
Patent Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6190128B1 (en) * | 1997-06-12 | 2001-02-20 | Mitsubishi Heavy Industries, Ltd. | Cooled moving blade for gas turbine |
US6328531B1 (en) * | 1998-08-05 | 2001-12-11 | Societe Nationale d'Etude et de Construction de Moteurs d'Aviation “SNECMA” | Cooled turbine blade |
US6062817A (en) * | 1998-11-06 | 2000-05-16 | General Electric Company | Apparatus and methods for cooling slot step elimination |
US20030138322A1 (en) * | 2002-01-23 | 2003-07-24 | Snecma Moteurs | Moving blade for a high pressure turbine, the blade having a trailing edge of improved thermal behavior |
US6851924B2 (en) * | 2002-09-27 | 2005-02-08 | Siemens Westinghouse Power Corporation | Crack-resistance vane segment member |
US20050249593A1 (en) * | 2004-01-14 | 2005-11-10 | Snecma Moteurs | Cooling air evacuation slots of turbine blades |
US7249933B2 (en) * | 2005-01-10 | 2007-07-31 | General Electric Company | Funnel fillet turbine stage |
US7371046B2 (en) * | 2005-06-06 | 2008-05-13 | General Electric Company | Turbine airfoil with variable and compound fillet |
US7581924B2 (en) * | 2006-07-27 | 2009-09-01 | Siemens Energy, Inc. | Turbine vanes with airfoil-proximate cooling seam |
Cited By (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9045987B2 (en) | 2012-06-15 | 2015-06-02 | United Technologies Corporation | Cooling for a turbine airfoil trailing edge |
EP2868867A1 (de) * | 2013-10-29 | 2015-05-06 | Siemens Aktiengesellschaft | Turbinenschaufel |
WO2015063021A1 (de) * | 2013-10-29 | 2015-05-07 | Siemens Aktiengesellschaft | Turbinenschaufel mit einer mittigen ausblasung an der hinterkante |
CN105683505A (zh) * | 2013-10-29 | 2016-06-15 | 西门子股份公司 | 在后缘具有中央吹出部的涡轮叶片或轮叶 |
US10487669B2 (en) | 2013-10-29 | 2019-11-26 | Siemens Aktiengesellschaft | Turbine blade with a central blowout at the trailing edge |
WO2018031032A1 (en) * | 2016-08-12 | 2018-02-15 | Siemens Aktiengesellschaft | Blade for gas turbine engine |
Also Published As
Publication number | Publication date |
---|---|
DE112009002628T5 (de) | 2012-01-26 |
WO2010037659A1 (de) | 2010-04-08 |
CH699601A1 (de) | 2010-03-31 |
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Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: ALSTOM TECHNOLOGY LTD, SWITZERLAND Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:NAGLER, CHRISTOPH;KREISELMAIER, ERICH;ANGUISOLA, JOSE MCFEAT;AND OTHERS;SIGNING DATES FROM 20110601 TO 20110606;REEL/FRAME:026423/0875 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |