EP1881157B1 - Serpentinenartige Mikrokanäle zur lokalen Wärmeabfuhr - Google Patents

Serpentinenartige Mikrokanäle zur lokalen Wärmeabfuhr Download PDF

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Publication number
EP1881157B1
EP1881157B1 EP20070252841 EP07252841A EP1881157B1 EP 1881157 B1 EP1881157 B1 EP 1881157B1 EP 20070252841 EP20070252841 EP 20070252841 EP 07252841 A EP07252841 A EP 07252841A EP 1881157 B1 EP1881157 B1 EP 1881157B1
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EP
European Patent Office
Prior art keywords
cooling
turbine engine
engine component
component according
circuit
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.)
Active
Application number
EP20070252841
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English (en)
French (fr)
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EP1881157A1 (de
Inventor
Francisco J. Cunha
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.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
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Publication date
Priority claimed from US11/489,155 external-priority patent/US7513744B2/en
Priority claimed from US11/494,831 external-priority patent/US7581928B1/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP1881157A1 publication Critical patent/EP1881157A1/de
Application granted granted Critical
Publication of EP1881157B1 publication Critical patent/EP1881157B1/de
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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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.
  • the overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design.
  • the ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil.
  • the opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away.
  • existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
  • Fig. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point 10 related to a new advanced serpentine microcircuit shown in FIGS. 2a - 2c .
  • This serpentine microcircuit includes a pressure side serpentine circuit 20 and a suction side serpentine circuit 22 embedded in the airfoil walls 24 and 26.
  • FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2a - 2c embedded in the airfoils walls.
  • FIGS. 4A and 4B There are however field problems that can be addressed efficiently with peripheral microcircuit designs.
  • FIG. 4A the streamlines of the gas path close to the external surface of the airfoil illustrate four different regions in which the gas flow changes direction or migration: a tip region, two midsection regions, and a root region. In between the tip and the upper mid region, the flow transitions through a pseudo stagnation point(s). The momentum of the external gas seems to decelerate in such a way as to impose a local thermal load to the part. This manifests itself by regions where the propensity for erosion and oxidation increase in the airfoil surface. The superposition of FIG.
  • 4B illustrates the local coincidence between the pseudo-stagnation region and the blade distress in the part surface.
  • the upper and lower region also converge onto one another, but even though the space between streamlines decreases, the flow seems to accelerate and there is no pseudo-stagnation regions.
  • a mild manifestation of the same tip-to-mid phenomena seems to initiate in the transition region between the mid-to-root regions. It is therefore necessary to tailor the peripheral microcircuit in such a manner as to address these local high thermal load regions.
  • a turbine engine component is provided with improved cooling as claimed in claim 1.
  • the two peripheral cooling arrangements include a peripheral pressure side microcircuit 100 which is incorporated or embedded within the wall forming the pressure side of an airfoil portion 104 and a suction side microcircuit 120 which is incorporated or embedded within the wall forming the suction side of the airfoil portion 104.
  • the pressure side peripheral microcircuit 100 is shown.
  • the first leg 102 has an inlet 103 which receives cooling fluid from a source (not shown).
  • the leg 102 provides a flow of cooling fluid which quenches the hot spot in the tip-to-mid region of the airfoil portion 104 shown in FIG. 4B .
  • the cooling fluid within the leg 102 proceeds around a 180 degree bend 106 which is supplemented with a plurality of film holes 108, preferably three film holes.
  • the film holes 108 ensure flow acceleration through the bend 106 to a second downstream leg 110 which ends below the platform 112 of the turbine engine component 90 in an exit 164. Cooling fluid from the leg 110 is fed into an internal trailing edge circuit 114 to be discussed hereinafter via the exit 164 where it is used to further cool the airfoil portion 104.
  • the circuit 120 has a first leg 122 which communicates with a source (not shown) of cooling fluid. In the first leg 122, the cooling flow convects heat away from the suction side. Since the circuit 120 has no film holes, effective cooling may not be done past the external gage point of the airfoil portion 104 where any film cooling would provide high aerodynamic penalties due to mixing. Thus, the circuit 120 is used to feed cooling fluid to a leading edge microcircuit 124 which wraps around the leading edge 126 of the airfoil portion 104. The circuit 120 feeds or supplies cooling fluid to the leading edge wrap around circuit 124 through a plurality of wall cross over holes 128.
  • the circuit 120 has a bend 130 and a second leg 132.
  • the holes 128 are preferably located in the vicinity of the bend 130 and the second leg 132.
  • the second leg 132 may also communicate with the wrap around circuit 124 via a passageway 134.
  • several holes 136 are located in the leading edge and are used to cool the leading edge of the airfoil portion 104.
  • the microcircuit 124 is provided with a plurality of film holes 138 for creating a film of cooling fluid over the pressure side of the airfoil portion.
  • the main body internal cooling circuits which include a leading edge internal cooling circuit 150 and the trailing edge internal cooling circuit 114.
  • the leading edge internal cooling circuit 150 communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet 151 and has one or more film cooling holes 152 adjacent the tip 154 of the airfoil portion 104 to provide tip cooling.
  • the circuit 150 also has a plurality of cross-over holes 156 for supplying cooling fluid to the leading edge microcircuit 124.
  • the trailing edge internal circuit 114 also communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet 157 and has one or more film cooling holes 158 adjacent the tip 154 to provide tip cooling.
  • the circuit 114 also has a plurality of cross-over holes 160 for communicating with a trailing edge cooling circuit 162 for cooling the trailing edge of the airfoil portion 104.
  • the trailing edge internal circuit 114 also receives cooling fluid from the peripheral pressure side microcircuit 100 via the exit 164.
  • Each of the leading edge internal circuit 150 and the trailing edge internal circuit 114 may be provided with a plurality of film cooling holes 170 and 172 respectively to form cooling films over the pressure and suction sides of the airfoil portion 104.
  • the airfoil portion of a turbine engine component may be very effectively convectively cooled.
  • the cooling flow is returned to the trailing edge internal circuit for further cooling of the airfoil.
  • the suction side circuit the leading edge of the airfoil is cooled first before discharging in pressure side film. This effective use of coolant allows for positive effects on cycle thermodynamic efficiency, turbine efficiency, rotor inlet temperature impacts, and specific fuel consumption.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (16)

  1. Turbinenmaschinenkomponente (90) aufweisend:
    einen Strömungsprofilbereich (104) aufweisend eine Druckseite und eine Zugseite;
    einen ersten Kühlkreislauf (100), der in der Wand, die die Druckseite des Strömungsprofilbereichs (104) bildet, eingelassen ist um den Druckbereich des Strömungsprofilbereichs (104) zu kühlen; und gekennzeichnet durch
    einen zweiten Kühlkreislauf (120), der in der Wand, die die Sogseite des Strömungsprofilbereichs (104) bildet, eingelassen ist um den Sogbereich des Strömungsprofilbereichs (104) zu kühlen und um Kühlfluid einem Mittel zum Herstellen eines Kühlungsfilms über der Druckseite zur Verfügung zu stellen, wobei das Mittel zum Herstellen eines Kühlungsfilms über der Druckseite einen Kühlkreislauf (124) aufweist, der um eine Vorderkante (126) des Strömungsprofilbereichs (104) herumgeführt ist.
  2. Turbinenmaschinenkomponente nach Anspruch 1, wobei weiterhin der Kühlkreislauf (124), der um die Vorderkante (125) herumgeführt ist und mit einem ersten Satz an Filmlöchern (138) zum Kühlen der Vorderkante ausgestattet ist.
  3. Turbinenmaschinenkomponente nach Anspruch 2, wobei weiterhin der Kühlkreislauf (124), der um die Vorderkante (126) herumgeführt ist und mit einem zweiten Satz an Filmlöchern (138) zum Kühlen der Druckseite des Strömungsprofilbereichs (104) ausgestattet ist.
  4. Turbinenmaschinenkomponente nach einem der vorherigen Ansprüche, weiterhin aufweisend einen inneren Vorderkantenkreislauf (150) und einen inneren Hinterkantenkreislauf (114).
  5. Turbinenmaschinenkomponente nach Anspruch 4, wobei der erste Kühlkreislauf (100) einen Ausgang (164) aufweist, der Kühlfluid zu dem inneren Hinterkantenkreislauf (114) liefert.
  6. Turbinenmaschinenkomponente nach Anspruch 5, wobei der erste Kühlkreislauf (100) eine erste Strecke (102), eine zweite Strecke (110) und eine Biegung (106) zwischen der ersten Strecke (102) und der zweiten Strecke (110) aufweist.
  7. Turbinenmaschinenkomponente nach Anspruch 6, wobei die zweite Strecke (110) in dem Ausgang (164) endet.
  8. Turbinenmaschinenkomponente nach Ansprüchen 6 oder 7, weiterhin aufweisend Mittel, zum Sicherstellen von Strömungsbeschleunigung durch die Biegung (106).
  9. Turbinenmaschinenkomponente nach Anspruch 8, wobei die Mittel zur Strömungsbeschleunigung eine Mehrzahl an Löchern (108) aufweisen.
  10. Turbinenmaschinenkomponente nach einem der Ansprüche 4 bis 9, wobei jeder der inneren Kreisläufe (150, 114) eine Mehrzahl an Filmlöchern zum Ausbilden einer Strömung des Kühlfluids über der Druckseite und über der Sogseite ausweist.
  11. Turbinenmaschinenkomponente nach einem der Ansprüche 4 bis 10, wobei der innere Vorderkantenkreislauf (150) eine Mehrzahl an Durchgangslöchern (156) aufweist, um das Fluid zu dem Vorderkantenkühlkreislauf (124) zu leiten.
  12. Turbinenmaschinenkomponente nach einem der vorherigen Ansprüche 4 bis 11, wobei der innere Hinterkantenkreislauf (114) eine Mehrzahl an Übergangslöchern (160) aufweist, um das Fluid zu dem Hinterkantenkühlkreislauf (162) zu leiten.
  13. Turbinenmaschinenkomponente nach einem der vorherigen Ansprüche 4 bis 12, wobei sowohl der innere Vorderkantenkreislauf als auch der innere Hinterkantenkreislauf (150, 114) Mittel zum Kühlen einer Spitze (154) des Strömungsprofilbereichs (104) aufweist.
  14. Turbinenmaschinenkomponente nach einem der vorherigen Ansprüche, wobei der zweite Kühlkreislauf (120) eine erste Strecke (122), eine zweite Strecke (132) und eine Biegung (130) zwischen der ersten Strecke (120) und der zweiten Strecke (132) aufweist.
  15. Turbinenmaschinenkomponente nach Anspruch 14, wobei der zweite Kühlkreislauf eine Mehrzahl an Übergangslöchern (128) aufweist, um ein Kühlfluid als Mittel zum Erzeugen eines Kühlungsfilms über der Druckseite bereitszustellen.
  16. Turbinenmaschinenkomponente nach Anspruch 14 oder 15, wobei die zweite Strecke (132) mit den Mitteln in Verbindung steht, um eine Kühlschicht über der Druckseite zu erzeugen.
EP20070252841 2006-07-18 2007-07-18 Serpentinenartige Mikrokanäle zur lokalen Wärmeabfuhr Active EP1881157B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US11/489,155 US7513744B2 (en) 2006-07-18 2006-07-18 Microcircuit cooling and tip blowing
US11/494,831 US7581928B1 (en) 2006-07-28 2006-07-28 Serpentine microcircuits for hot gas migration

Publications (2)

Publication Number Publication Date
EP1881157A1 EP1881157A1 (de) 2008-01-23
EP1881157B1 true EP1881157B1 (de) 2014-02-12

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Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8753083B2 (en) * 2011-01-14 2014-06-17 General Electric Company Curved cooling passages for a turbine component
EP2752554A1 (de) * 2013-01-03 2014-07-09 Siemens Aktiengesellschaft Schaufel für eine Turbomaschine

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5813835A (en) 1991-08-19 1998-09-29 The United States Of America As Represented By The Secretary Of The Air Force Air-cooled turbine blade
US5931638A (en) * 1997-08-07 1999-08-03 United Technologies Corporation Turbomachinery airfoil with optimized heat transfer
US6254334B1 (en) * 1999-10-05 2001-07-03 United Technologies Corporation Method and apparatus for cooling a wall within a gas turbine engine
GB0114503D0 (en) 2001-06-14 2001-08-08 Rolls Royce Plc Air cooled aerofoil
US6981846B2 (en) * 2003-03-12 2006-01-03 Florida Turbine Technologies, Inc. Vortex cooling of turbine blades
US7097426B2 (en) 2004-04-08 2006-08-29 General Electric Company Cascade impingement cooled airfoil
US7011502B2 (en) * 2004-04-15 2006-03-14 General Electric Company Thermal shield turbine airfoil

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