Serpentine microcircuits for local heat removal

Description

BACKGROUND

(1) Field of the Invention

[0001] The present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.

(2) Prior Art

[0002] 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. In general, 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.

[0003] 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.

[0004] The Table I below provides the operational parameters used to plot the design point in the durability map.

TABLE I

Operational Parameters for serpentine microcircuit
beta	2.898
Tg	2581 [F]
Tc	1365 [F]
Tm	2050 [F]
Tm_bulk	1709 [F]
Phi_loc	0.437
Phi_bulk	0.717
Tco	1640 [F]
Tci	1090 [F]
eta_c_loc	0.573
eta_f	0.296
Cooling Flow	3.503%
Total WAE	10.8

Legend for Table I Beta = heat load Phi_loc = local cooling effectiveness Phi_bulk = bulk cooling effectiveness Eta_c_loc = local cooling efficiency Eta_f = film effectiveness Tg = gas temperature Tc = coolant temperature Tm = metal temperature Tm_bulk = bulk metal temperature Tco = exit coolant temperature Tci = inlet coolant temperature WAE = compressor engine flow, pps

[0005] It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow. FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2a - 2c embedded in the airfoils walls.

[0006] There are however field problems that can be addressed efficiently with peripheral microcircuit designs. One such field problem is illustrated in FIGS. 4A and 4B. In 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. In the mid region, 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.

[0007] US 5,813,835, EP 1 267 038 and EP 1 584 790 all describe air cooled airfoils.

SUMMARY OF THE INVENTION

[0008] In accordance with the present invention, a turbine engine component is provided with improved cooling as claimed in claim 1.

[0009] Other details of the serpentine microcircuits for hot gas migration of the present invention, as well as other advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.

BRIEF DESCRIPTION OF THE DRAWINGS

[0010] 

FIG. 1 is a graph showing cooling effectiveness versus film effectiveness for a turbine engine component;

FIG. 2A shows an airfoil portion of a turbine engine component having a pressure side cooling microcircuit embedded in the pressure side wall and a suction side cooling microcircuit embedded in the suction side wall;

FIG. 2B is a schematic representation of a pressure side cooling microcircuit used in the airfoil portion of FIG. 2A;

FIG. 2C is a schematic representation of a suction side cooling microcircuit used in the airfoil portion of FIG. 2A;

FIG. 3 illustrates the cooling flow distribution for a turbine engine component with serpentine microcircuits embedded in the airfoil walls;

FIG. 4A is a schematic representation illustrating the pressure side distress on an airfoil surface;

FIG. 4B is a schematic representation of the local coincidence between the pseudo-stagnation region and the blade distress;

FIG. 5 is a schematic representation of a peripheral pressure side cooling circuit;

FIG. 6 is a schematic representation of a peripheral suction side cooling circuit; and

FIG. 7 is a schematic representation of main body internal cooling circuits.

DETAILED DESCRIPTION OF THE PREFERED EMBODIMENT(S)

[0011] Referring now to FIGS. 5 and 6, there are depicted two peripheral cooling arrangements which may be used to address local increases in the airfoil thermal load of a turbine engine component 90 such as a turbine blade. 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.

[0012] In FIG. 5, the pressure side peripheral microcircuit 100 is shown. In this circuit, 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.

[0013] Referring now to FIG. 6, there is shown a peripheral suction side microcircuit 120. 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. As can be seen from FIG. 6, 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. As the microcircuit 124 wraps around the leading edge, several holes 136 are located in the leading edge and are used to cool the leading edge of the airfoil portion 104. Further, 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.

[0014] Referring now to FIG. 7, there is shown 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.

[0015] 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. As can be seen from FIG. 7, the trailing edge internal circuit 114 also receives cooling fluid from the peripheral pressure side microcircuit 100 via the exit 164.

[0016] 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.

[0017] Using the pressure and suction side cooling circuits of the present invention, the airfoil portion of a turbine engine component may be very effectively convectively cooled. Using the pressure side circuit, the cooling flow is returned to the trailing edge internal circuit for further cooling of the airfoil. Using 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.

Claims

1. A turbine engine component (90) comprising:

an airfoil portion (104) having a pressure side and a suction side;

a first cooling circuit (100) embedded within the wall forming the pressure side of the airfoil portion (104) for cooling said pressure side of said airfoil portion (104); and characterised by

a second cooling circuit (120) embedded within the wall forming the suction side of the airfoil portion (104) for cooling said suction side of said airfoil portion (104) and for supplying cooling fluid to means for creating a cooling film over said pressure side, wherein said means for creating a cooling film over said pressure side comprises a cooling circuit (124) wrapped around a leading edge (126) of said airfoil portion (104).

2. The turbine engine component according to claim 1, further comprising said cooling circuit (124) wrapped around said leading edge (126) having a first set of film holes (138) for cooling said leading edge.

3. The turbine engine component according to claim 2 further comprising said cooling circuit (124) wrapped around said leading edge (126) having a second set of film holes (138) for cooling said pressure side of said airfoil portion (104).

4. The turbine engine component according to any preceding claim, further comprising a leading edge internal circuit (150) and a trailing edge internal circuit (114).

5. The turbine engine component according to claim 4, wherein said first cooling circuit (100) has an exit (164) which delivers cooling fluid to said trailing edge internal circuit (114).

6. The turbine engine component according to claim 5, wherein said first cooling circuit (100) has a first leg (102), a second leg (110), and a bend (106) between said first leg (102) and said second leg (110).

7. The turbine engine component according to claim 6, wherein said second leg (110) terminates in said exit (164).

8. The turbine engine component according to claim 6 or 7, further comprising means for ensuring flow acceleration through the bend (106).

9. The turbine engine component according to claim 8, wherein said flow acceleration ensuring means comprises a plurality of holes (108).

10. The turbine engine component according to any of claims 4 to 9, wherein each of said internal circuits (150, 114) has a plurality of film holes for creating a flow of cooling fluid over said pressure side and said suction side.

11. The turbine engine component according to any of claims 4 to 10, wherein said leading edge internal circuit (150) has a plurality of cross-over holes (156) for supplying fluid to a leading edge cooling circuit (124).

12. The turbine engine component according to any of claims 4 to 11, wherein said trailing edge internal circuit (114) has a plurality of cross-over holes (160) for supplying fluid to a trailing edge cooling circuit (162).

13. The turbine engine component according to any of claims 4 to 12, wherein each of said leading edge and said trailing edge internal circuits (150, 114) has means for cooling a tip (154) of said airfoil portion (104).

14. The turbine engine component according to any preceding claim, wherein said second cooling circuit (120) has a first leg (122), a second leg (132), and a bend (130) between said first leg (120) and said second leg (132).

15. The turbine engine component according to claim 14, wherein said second cooling circuit (120) has a plurality of cross-over holes (128) for supplying cooling fluid to said means for creating a cooling film over said pressure side.

16. The turbine engine component according to claim 14 or 15, wherein said second leg (132) communicates with said means for creating a cooling film over said pressure side.

Ansprüche

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.

Revendications

1. Composant de moteur à turbine (90) comprenant :

une partie formant profil aérodynamique (104) ayant un côté refoulement et un côté aspiration ;

un premier circuit de refroidissement (100) incorporé à l'intérieur de la paroi formant le côté refoulement de la partie formant profil aérodynamique (104) pour refroidir ledit côté refoulement de ladite partie formant profil aérodynamique (104) ; et caractérisé par

un second circuit de refroidissement (120) incorporé à l'intérieur de la paroi formant le côté aspiration de la partie formant profil aérodynamique (104) pour refroidir ledit côté aspiration de ladite partie formant profil aérodynamique (104) et pour fournir du fluide de refroidissement à un moyen pour créer un film de refroidissement sur ledit côté refoulement, dans lequel ledit moyen pour créer un film de refroidissement sur ledit côté refoulement comprend un circuit de refroidissement (124) enroulé autour d'un bord d'attaque (126) de ladite partie formant profil aérodynamique (104).

2. Composant de moteur à turbine selon la revendication 1, comprenant en outre ledit circuit de refroidissement (124) enroulé autour dudit bord d'attaque (126) ayant un premier ensemble de trous de refroidissement par convection (138) pour refroidir ledit bord d'attaque.

3. Composant de moteur à turbine selon la revendication 2, comprenant en outre ledit circuit de refroidissement (124) enroulé autour dudit bord d'attaque (126) ayant un second ensemble de trous de refroidissement par convection (138) pour refroidir ledit côté refoulement de ladite partie formant profil aérodynamique (104).

4. Composant de moteur à turbine selon l'une quelconque des revendications précédentes, comprenant en outre un circuit interne de bord d'attaque (150) et un circuit interne de bord de fuite (114).

5. Composant de moteur à turbine selon la revendication 4, dans lequel ledit premier circuit de refroidissement (100) a une sortie (164) qui délivre du fluide de refroidissement audit circuit interne de bord de fuite (114).

6. Composant de moteur à turbine selon la revendication 5, dans lequel ledit premier circuit de refroidissement (100) a une première branche (102), une seconde branche (110) et une courbure (106) entre ladite première branche (102) et ladite seconde branche (110).

7. Composant de moteur à turbine selon la revendication 6, dans lequel ladite seconde branche (110) se termine dans ladite sortie (164).

8. Composant de moteur à turbine selon la revendication 6 ou 7, comprenant en outre un moyen pour assurer une accélération de flux à travers la courbure (106).

9. Composant de moteur à turbine selon la revendication 8, dans lequel ledit moyen assurant l'accélération de flux comprend une pluralité de trous (108).

10. Composant de moteur à turbine selon l'une quelconque des revendications 4 à 9, dans lequel chacun desdits circuits internes (150, 114) a une pluralité de trous de refroidissement par convection pour créer un flux de fluide de refroidissement sur ledit côté refoulement et ledit côté aspiration.

11. Composant de moteur à turbine selon l'une quelconque des revendications 4 à 10, dans lequel ledit circuit interne de bord d'attaque (150) a une pluralité de trous de liaison (156) pour fournir du fluide à un circuit de refroidissement de bord d'attaque (124).

12. Composant de moteur à turbine selon l'une quelconque des revendications 4 à 11, dans lequel ledit circuit interne de bord de fuite (114) a une pluralité de trous de liaison (160) pour fournir du fluide à un circuit de refroidissement de bord de fuite (162).

13. Composant de moteur à turbine selon l'une quelconque des revendications 4 à 12, dans lequel chacun desdits circuits internes de bord d'attaque et de bord de fuite (150, 114) a un moyen pour refroidir une extrémité (154) de ladite partie dormant profil aérodynamique (104).

14. Composant de moteur à turbine selon l'une quelconque des revendications précédentes, dans lequel ledit second circuit de refroidissement (120) a une première branche (122), une seconde branche (132), et une courbure (130) entre ladite première branche (120) et ladite seconde branche (132).

15. Composant de moteur à turbine selon la revendication 14, dans lequel ledit second circuit de refroidissement (120) a une pluralité de trous de liaison (128) pour fournir du fluide de refroidissement audit moyen pour créer un film de refroidissement sur ledit côté refoulement.

16. Composant de moteur à turbine selon la revendication 14 ou 15, dans lequel ladite seconde branche (132) communique avec ledit moyen pour créer un film de refroidissement sur ledit côté refoulement.

Drawing