EP0207799A2 - Kühlkanäle für die Schaufeln einer Gasturbine - Google Patents
Kühlkanäle für die Schaufeln einer Gasturbine Download PDFInfo
- Publication number
- EP0207799A2 EP0207799A2 EP86305182A EP86305182A EP0207799A2 EP 0207799 A2 EP0207799 A2 EP 0207799A2 EP 86305182 A EP86305182 A EP 86305182A EP 86305182 A EP86305182 A EP 86305182A EP 0207799 A2 EP0207799 A2 EP 0207799A2
- Authority
- EP
- European Patent Office
- Prior art keywords
- flow area
- blade
- section
- coolant
- hub
- 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.)
- Withdrawn
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
- F05D2250/00—Geometry
- F05D2250/20—Three-dimensional
- F05D2250/23—Three-dimensional prismatic
- F05D2250/231—Three-dimensional prismatic cylindrical
-
- 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
- F05D2250/23—Three-dimensional prismatic
- F05D2250/232—Three-dimensional prismatic conical
Definitions
- the present invention relates to rotor blades in a combustion turbine and more particularly to coolant systems therefor.
- Rotating turbine blades are frequently cooled with air flowing radially outward through a plurality of holes which pass from the blade root to the blade tip.
- the holes typically have either a constant diameter along the airfoil and root portions of the holes or a first constant diameter along the airfoil portion and a second constant diameter along the root portion of the holes.
- the diameter of the hole along the root portion typically being larger to avoid pressure loss in a region that does not require appreciable cooling.
- the airfoil portions of the coolant holes need to be relatively small in cross section to produce the high coolant velocity and heat transfer coefficient required there.
- the critical design region of the blade for both stress and cooling is the center span portion of the blade, and the hole diameter, the number of holes and the coolant flow are normally set by design considerations. Since the coolant heats up appreciably as it flows outwardly along the blade, having received heat from the hot blade path gas, the coolant temperature at the center span region becomes considerably higher than the coolant temperature near the blade hub.
- the lower coolant temperature near the blade hub tends to cool the blade to a lower temperature than is needed from the standpoint of stress design.
- Overcoo'ling near the blade hub means the coolant absorbs more heat than necessary from the hub region which results in increased coolant temperature in the center span region and thus higher coolant flows and/or higher metal temperatures than if the overcooling did not occur.
- overcooling since reduced centrifugal stress usually more than offsets the coolant heatup in the blade tip region, overcooling also tends to occur there.
- An important consequence of overcooling in both the hub and tip regions is a higher level of pressure loss than would be encountered if the coolant flow per unit of flow area in overcooled regions were reduced to produce a heat transfer level matching the cooling requirements.
- Reduction of pressure loss in overcooled regions would make possible higher cooling flow per unit of flow area in the mid-span region in order to provide increased cooling for a given supply pressure at the blade root. This can be translated into a design for higher turbine inlet temperature or one with reduced cooling flow for given turbine inlet temperatures.
- cores are typically used to form the radial coolant holes in a blade. Structural weakness of the cores has resulted in core breakage and blade scrapping more frequently than desirable thereby adding to per unit manufacturing costs of blades.
- the present invention resides in a combustion turbine rotor blade having a root portion and an airfoil portion extending therefrom, said airfoil portion having a plurality of coolant holes extending from said root portion along its span to its tip to provide flow of coolant therethrough from said root portion, each of said coolant holes having an inner hub portion and a second portion extending outwardly from said hub portion, characterized by said hub and second hole portions having relative flow areas which cause the coolant to have reduced flow per unit of flow area in said hub hole portion relative to the coolant flow per unit of flow area through said second hole portion over a mid-span region of said blade thereby avoiding hub blade region overcooling.
- FIG. 1 a combustion turbine blade 10 having a root portion 12 and an airfoil portion 14 arranged in accordance with the principles of the invention.
- Coolant holes or channels 16 extend radially outwardly along the span of the blade. Air flow supplied from complementary coolant holes in the blade root and directed outwardly along the coolant holes 16 cools the blade airfoil portion 14 and its surface which is exposed to the hot blade path gas.
- a center span portion 18 of the blade airfoil is the critical design region for which the number of holes, the hole diameter and the coolant flow are set to meet its cooling needs.
- the total blade coolant system structure is then designed to support the critical region needs.
- the temperature of the coolant in the holes 16 is lower in a blade hub region 20 than in the mid-span region 18 because the coolant heats up as it flows outwardly along the blade.
- the lower temperature coolant tends to overcool the blade hub 20 resulting in coolant temperature in the mid-span region 18 being higher than it would be if overcooling did not occur.
- coolant flow and/or metal temperature is higher than it could be if coolant temperature rise were reduced with the elimination of overcooling.
- blade hub overcooling is substantially reduced or eliminated to permit lower coolant flow and/or lower metal temperature resulting in greater cooling and turbine efficiency.
- the coolant hole 16 includes a hub portion 22 and an outer portion 24 which extends from the hub portion 22 through the blade mid-span region to the blade tip 26.
- the hub portion 22 of the coolant hole 16 is tapered from a first diameter at its inlet end to a smaller diameter where it joins the outer hole portion 24.
- the diameter of the outer hole portion 24 is substantially constant along its length.
- the hole flow area in the hub region 20 is increased relative to the hole flow area in the mid-span region 18.
- Reduced coolant flow per unit of flow area in the hub region 20 reduces the hub region heat transfer coefficient and the amount of cooling.
- Turbine and blade design parameters determine the amount and length of taper to reduce or substantially eliminate overcooling in the hub region 20.
- Tapering of the coolant holes also reduces pressure loss in the coolant holes. Accordingly, higher flow per unit of flow area is obtained in the mid-span. region to obtain lower metal temperature for a given supply pressure at the blade root. Spanwise hole cooling technology can thus be employed with higher turbine inlet temperature levels. The higher flow per unit of flow area in the design section can also be used to reduce cooling flow for a given level of turbine inlet temperatures.
- a coolant hole 16a includes a tapered hub portion 22a, a constant diameter mid-span portion 24a and a tip portion 25a which is tapered outwardly in the direction of coolant flow.
- the constant diameter mid-span portion 24a can be eliminated from the hole 16a so that the hole 16a is formed by the oppositely tapered portions 22a and 25a in end-to- end relation.
- the junction point of the hole portions 22a and 25a in this alternative would be determined by stress and heat transfer considerations.
- the tapered hole portions be provided with continuous tapering, either linear as shown or nonlinear as warranted by design considerations.
- the coolant holes have a circular cross-section but the invention can be implemented with non-circular cross- sections.
- FIG 4 there is shown another embodiment of the invention in which a coolant hole 16b has a stepped configuration to provide a varying cross-section along its length for substantially reduced blade region overcooling.
- the coolant hole 16b includes a hub portion 22b having an inner section 22bl having a first diameter and an outer section 22b2 having a lesser diameter.
- a mid-span section 24b having a further reduced diameter adjoins the outer hub section 22b2.
- a tip portion 25b of the coolant hole 16b has a first section 25bl and a second section 25b2 with successively greater diameters.
- tapered blade coolant holes provides a further advantage by permitting the blades to be manufactured by casting processes using cores for coolant hole formation.
- tapered cores can be employed to form the coolant holes 16 and 16a, and tapered cores with larger diameter in the hub and tip region are characteristically stronger than cores conventionally used to form constant diameter coolant holes in the airfoil. With stronger cores, there is reduced core breakage, reduced blade scrapping and reduced manufacturing costs.
- the stronger tapered cores permit manufacture of smaller hole diameters in the mid-span portion of the coolant holes. In turn, improved blade cooling and reduced blade coolant flow can be realized.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US75165785A | 1985-07-03 | 1985-07-03 | |
US751657 | 1985-07-03 |
Publications (2)
Publication Number | Publication Date |
---|---|
EP0207799A2 true EP0207799A2 (de) | 1987-01-07 |
EP0207799A3 EP0207799A3 (de) | 1988-09-14 |
Family
ID=25022941
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP86305182A Withdrawn EP0207799A3 (de) | 1985-07-03 | 1986-07-03 | Kühlkanäle für die Schaufeln einer Gasturbine |
Country Status (5)
Country | Link |
---|---|
EP (1) | EP0207799A3 (de) |
JP (1) | JPS6210402A (de) |
CN (1) | CN86104500A (de) |
CA (1) | CA1262868A (de) |
IE (1) | IE861475L (de) |
Cited By (23)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0319758A1 (de) * | 1987-12-08 | 1989-06-14 | General Electric Company | Diffusionskühlung einer Schaufelspitze |
EP0550184A1 (de) * | 1991-12-30 | 1993-07-07 | General Electric Company | Kühlkanäle mit Turbulenzpromotoren für Gasturbinenschaufeln |
US6539627B2 (en) | 2000-01-19 | 2003-04-01 | General Electric Company | Method of making turbulated cooling holes |
EP1422383A2 (de) * | 2002-11-20 | 2004-05-26 | Mitsubishi Heavy Industries, Ltd. | Kühlung einer Gasturbinenschaufel |
WO2004057157A1 (en) * | 2002-12-21 | 2004-07-08 | John Macdonald | Turbine blade |
EP1541805A1 (de) * | 2003-12-12 | 2005-06-15 | General Electric Company | Schaufel mit Kühllöchern |
EP1820937A2 (de) | 2006-02-15 | 2007-08-22 | United Technologies Corporation | Turbinenschaufel mit Radial-Kühlkanälen |
US20100329862A1 (en) * | 2009-06-24 | 2010-12-30 | General Electric Company | Cooling Hole Exits for a Turbine Bucket Tip Shroud |
EP2354453A1 (de) * | 2010-02-02 | 2011-08-10 | Siemens Aktiengesellschaft | Turbinenmotorkomponente zur adaptiven Kühlung |
US8157527B2 (en) * | 2008-07-03 | 2012-04-17 | United Technologies Corporation | Airfoil with tapered radial cooling passage |
US20130052035A1 (en) * | 2011-08-24 | 2013-02-28 | General Electric Company | Axially cooled airfoil |
US8506251B2 (en) | 2010-03-03 | 2013-08-13 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade |
US8511992B2 (en) * | 2008-01-22 | 2013-08-20 | United Technologies Corporation | Minimization of fouling and fluid losses in turbine airfoils |
EP2743454A1 (de) * | 2012-12-11 | 2014-06-18 | General Electric Company | Turbinenkomponente mit Kühlbohrungen mit variierendem Durchmesser |
US9200526B2 (en) | 2010-12-21 | 2015-12-01 | Kabushiki Kaisha Toshiba | Transition piece between combustor liner and gas turbine |
WO2017121689A1 (en) | 2016-01-15 | 2017-07-20 | Siemens Aktiengesellschaft | Gas turbine aerofoil |
US10408079B2 (en) | 2015-02-18 | 2019-09-10 | Siemens Aktiengesellschaft | Forming cooling passages in thermal barrier coated, combustion turbine superalloy components |
US10686199B2 (en) | 2012-08-14 | 2020-06-16 | Loop Energy Inc. | Fuel cell flow channels and flow fields |
US10734661B2 (en) | 2012-08-14 | 2020-08-04 | Loop Energy Inc. | Fuel cell components, stacks and modular fuel cell systems |
US10930942B2 (en) | 2016-03-22 | 2021-02-23 | Loop Energy Inc. | Fuel cell flow field design for thermal management |
US11060195B2 (en) | 2012-08-14 | 2021-07-13 | Loop Energy Inc. | Reactant flow channels for electrolyzer applications |
CN113454322A (zh) * | 2019-03-29 | 2021-09-28 | 三菱动力株式会社 | 高温部件以及高温部件的制造方法 |
US20220170376A1 (en) * | 2019-06-05 | 2022-06-02 | Mitsubishi Power Ltd. | Turbine blade, manufacturing method for turbine blade, and gas turbine |
Families Citing this family (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE4430302A1 (de) * | 1994-08-26 | 1996-02-29 | Abb Management Ag | Prallgekühltes Wandteil |
CN1318735C (zh) * | 2005-12-26 | 2007-05-30 | 北京航空航天大学 | 一种适用于燃气涡轮发动机的脉动冲击冷却叶片 |
JP2010053749A (ja) * | 2008-08-27 | 2010-03-11 | Mitsubishi Heavy Ind Ltd | タービン用翼 |
US8201621B2 (en) * | 2008-12-08 | 2012-06-19 | General Electric Company | Heat exchanging hollow passages with helicoidal grooves |
WO2011108164A1 (ja) * | 2010-03-03 | 2011-09-09 | 三菱重工業株式会社 | ガスタービンの動翼およびその製造方法ならびに動翼を用いたガスタービン |
US8727724B2 (en) * | 2010-04-12 | 2014-05-20 | General Electric Company | Turbine bucket having a radial cooling hole |
US9181819B2 (en) * | 2010-06-11 | 2015-11-10 | Siemens Energy, Inc. | Component wall having diffusion sections for cooling in a turbine engine |
US10577943B2 (en) * | 2017-05-11 | 2020-03-03 | General Electric Company | Turbine engine airfoil insert |
CN110159357B (zh) * | 2019-06-04 | 2021-01-29 | 北京航空航天大学 | 一种提升被动安全的航空发动机涡轮叶片缩扩型供气通道 |
JP6637630B1 (ja) * | 2019-06-05 | 2020-01-29 | 三菱日立パワーシステムズ株式会社 | タービン翼およびタービン翼の製造方法並びにガスタービン |
Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3301528A (en) * | 1964-11-13 | 1967-01-31 | Rolls Royce | Aerofoil shaped blade for fluid flow machines |
Family Cites Families (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS4825103B1 (de) * | 1967-06-05 | 1973-07-26 | ||
JPS5520041A (en) * | 1978-07-29 | 1980-02-13 | Noto Denshi Kogyo Kk | Piezoelectric device |
JPS5669423A (en) * | 1979-11-09 | 1981-06-10 | Hitachi Ltd | Air-cooled blade of gas turbine |
-
1986
- 1986-06-04 IE IE147586A patent/IE861475L/xx unknown
- 1986-06-25 CA CA000512386A patent/CA1262868A/en not_active Expired
- 1986-06-30 JP JP15185286A patent/JPS6210402A/ja active Pending
- 1986-07-01 CN CN 86104500 patent/CN86104500A/zh active Pending
- 1986-07-03 EP EP86305182A patent/EP0207799A3/de not_active Withdrawn
Patent Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3301528A (en) * | 1964-11-13 | 1967-01-31 | Rolls Royce | Aerofoil shaped blade for fluid flow machines |
Cited By (37)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4893987A (en) * | 1987-12-08 | 1990-01-16 | General Electric Company | Diffusion-cooled blade tip cap |
EP0319758A1 (de) * | 1987-12-08 | 1989-06-14 | General Electric Company | Diffusionskühlung einer Schaufelspitze |
EP0550184A1 (de) * | 1991-12-30 | 1993-07-07 | General Electric Company | Kühlkanäle mit Turbulenzpromotoren für Gasturbinenschaufeln |
US5413463A (en) * | 1991-12-30 | 1995-05-09 | General Electric Company | Turbulated cooling passages in gas turbine buckets |
US6539627B2 (en) | 2000-01-19 | 2003-04-01 | General Electric Company | Method of making turbulated cooling holes |
US6824360B2 (en) | 2000-01-19 | 2004-11-30 | General Electric Company | Turbulated cooling holes |
EP1422383A3 (de) * | 2002-11-20 | 2006-05-31 | Mitsubishi Heavy Industries, Ltd. | Kühlung einer Gasturbinenschaufel |
EP1422383A2 (de) * | 2002-11-20 | 2004-05-26 | Mitsubishi Heavy Industries, Ltd. | Kühlung einer Gasturbinenschaufel |
US6994514B2 (en) * | 2002-11-20 | 2006-02-07 | Mitsubishi Heavy Industries, Ltd. | Turbine blade and gas turbine |
WO2004057157A1 (en) * | 2002-12-21 | 2004-07-08 | John Macdonald | Turbine blade |
US6997679B2 (en) * | 2003-12-12 | 2006-02-14 | General Electric Company | Airfoil cooling holes |
EP1541805A1 (de) * | 2003-12-12 | 2005-06-15 | General Electric Company | Schaufel mit Kühllöchern |
EP1820937A2 (de) | 2006-02-15 | 2007-08-22 | United Technologies Corporation | Turbinenschaufel mit Radial-Kühlkanälen |
EP1820937A3 (de) * | 2006-02-15 | 2011-02-23 | United Technologies Corporation | Turbinenschaufel mit Radial-Kühlkanälen |
US8511992B2 (en) * | 2008-01-22 | 2013-08-20 | United Technologies Corporation | Minimization of fouling and fluid losses in turbine airfoils |
US8157527B2 (en) * | 2008-07-03 | 2012-04-17 | United Technologies Corporation | Airfoil with tapered radial cooling passage |
US20100329862A1 (en) * | 2009-06-24 | 2010-12-30 | General Electric Company | Cooling Hole Exits for a Turbine Bucket Tip Shroud |
JP2011007181A (ja) * | 2009-06-24 | 2011-01-13 | General Electric Co <Ge> | タービンバケット先端シュラウド用の冷却孔出口 |
US8511990B2 (en) * | 2009-06-24 | 2013-08-20 | General Electric Company | Cooling hole exits for a turbine bucket tip shroud |
EP2354453A1 (de) * | 2010-02-02 | 2011-08-10 | Siemens Aktiengesellschaft | Turbinenmotorkomponente zur adaptiven Kühlung |
US8827646B2 (en) | 2010-03-03 | 2014-09-09 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade |
US8506251B2 (en) | 2010-03-03 | 2013-08-13 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade |
US9200526B2 (en) | 2010-12-21 | 2015-12-01 | Kabushiki Kaisha Toshiba | Transition piece between combustor liner and gas turbine |
US20130052035A1 (en) * | 2011-08-24 | 2013-02-28 | General Electric Company | Axially cooled airfoil |
US11060195B2 (en) | 2012-08-14 | 2021-07-13 | Loop Energy Inc. | Reactant flow channels for electrolyzer applications |
US10686199B2 (en) | 2012-08-14 | 2020-06-16 | Loop Energy Inc. | Fuel cell flow channels and flow fields |
US10734661B2 (en) | 2012-08-14 | 2020-08-04 | Loop Energy Inc. | Fuel cell components, stacks and modular fuel cell systems |
US11489175B2 (en) | 2012-08-14 | 2022-11-01 | Loop Energy Inc. | Fuel cell flow channels and flow fields |
EP2743454A1 (de) * | 2012-12-11 | 2014-06-18 | General Electric Company | Turbinenkomponente mit Kühlbohrungen mit variierendem Durchmesser |
US10408079B2 (en) | 2015-02-18 | 2019-09-10 | Siemens Aktiengesellschaft | Forming cooling passages in thermal barrier coated, combustion turbine superalloy components |
WO2017121689A1 (en) | 2016-01-15 | 2017-07-20 | Siemens Aktiengesellschaft | Gas turbine aerofoil |
US10930942B2 (en) | 2016-03-22 | 2021-02-23 | Loop Energy Inc. | Fuel cell flow field design for thermal management |
US11901591B2 (en) | 2016-03-22 | 2024-02-13 | Loop Energy Inc. | Fuel cell flow field design for thermal management |
CN113454322A (zh) * | 2019-03-29 | 2021-09-28 | 三菱动力株式会社 | 高温部件以及高温部件的制造方法 |
CN113454322B (zh) * | 2019-03-29 | 2023-10-31 | 三菱动力株式会社 | 高温部件以及高温部件的制造方法 |
US20220170376A1 (en) * | 2019-06-05 | 2022-06-02 | Mitsubishi Power Ltd. | Turbine blade, manufacturing method for turbine blade, and gas turbine |
US11905848B2 (en) * | 2019-06-05 | 2024-02-20 | Mitsubishi Heavy Industries, Ltd. | Turbine blade, manufacturing method for turbine blade, and gas turbine |
Also Published As
Publication number | Publication date |
---|---|
IE861475L (en) | 1987-01-03 |
CN86104500A (zh) | 1987-02-04 |
EP0207799A3 (de) | 1988-09-14 |
JPS6210402A (ja) | 1987-01-19 |
CA1262868A (en) | 1989-11-14 |
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Inventor name: HOLDEN, PAUL CLARENCE |