Leading edge cooling of a gas turbine component using staggered turbulator strips

Description

BACKGROUND

(1) Field of the Invention

[0001] The present invention relates to enhanced cooling of the leading edge of airfoil portions of turbine engine components using trip strips that are preferably staggered and that are wrapped around the nose of the leading edge cavity.

(2) Prior Art

[0002] Due to the extreme environment in which they are used, some turbine engine components, such as blades and vanes, are cooled. A variety of different cooling techniques have been employed. One such scheme is illustrated in FIG. 1 where there is shown an airfoil portion 10 of a turbine engine component 12. As can be seen from the figure, a radial flow leading edge cavity 14 is used to effect cooling of the leading edge region. Another exemplary prior art cooling arrangement is shown in US 6,068,445.

[0003] Despite the existence of such a cooling scheme, there remains a need for improving the cooling of the leading edge of the airfoil portions of turbine engine components.

SUMMARY OF THE INVENTION

[0004] Accordingly, it is an aim of the present invention to provide enhanced cooling for the leading edge of airfoil portions of turbine engine components.

[0005] The present invention provides a turbine engine component as set forth in claim 1.

[0006] Other details of the leading edge cooling using staggered trip strips 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

[0007] 

FIG. 1 illustrates a prior art turbine engine component having a radial flow leading edge cavity;

FIG. 2 illustrates a cross-section of a leading edge portion of an airfoil used in a turbine engine component having staggered and wrapped trip strips;

FIG. 3 illustrates the trip strips on the suction side of the leading edge portion;

FIG. 4 illustrates the trip strips on the pressure side of the leading edge portion;

FIG. 5 illustrates the placement of the leading edge of the staggered trip strips;

FIG. 6 is a three dimensional view of the leading edge trip strips; and

FIG. 7 illustrates the vortex generated in the leading edge cavity.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)

[0008] Referring now to the drawings, FIG. 2 illustrates the leading edge 30 of an airfoil portion 32 of a turbine engine component. As can be seen from this figure, the leading edge 30 has a leading edge cavity 34 in which a cooling fluid, such as engine bleed air, flows in a radial direction. The leading edge 30 also has a nose portion 36 and an external stagnation region 38.

[0009] It has been found that trip strips are desirable to provide adequate cooling of the leading edge, especially at the nose portion 36 of the airfoil portion 32 adjacent to the external stagnation region 38. The trip strip arrangement which will be discussed hereinafter provides high heat transfer to the leading edge 30 of the airfoil portion 32.

[0010] As shown in FIGS. 2, 5 and 6, a plurality of trip strips 40 are positioned on the pressure side 42 of the airfoil portion 32, while, as shown in FIGS. 2, 3, and 6, a plurality of trip strips 44 are placed on the suction side 46 of the airfoil portion 32. Referring now to FIGS. 2 and 6, the trip strips 40 on the pressure side 42 are wrapped around the leading edge nose portion 36. As the pressure side trip strips 40 wrap around the leading edge, the curvature of the leading edge nose portion 36 causes the trip strips 40 to be oriented more or less normal to the direction of flow 48 (see FIG. 6). As cooling air passes over the thus oriented trip strips 40, the flow is tripped and generates a large vortex 49 at the leading edge (see FIG. 7). This large vortex generates very high heat transfer coefficients at the leading edge nose 36.

[0011] Referring now to FIGS. 4 and 6, it can be seen that the trip strips 40 and the trip strips 44 are preferably staggered approximately one half pitch apart between the suction side 46 and the pressure side 42 of the airfoil portion 32. As shown in FIGS. 2 and 7, there is also a gap 47 between adjacent ones of the trip strips 40 and the trip strips 44. Each gap 47 is located along a parting line 70 of the airfoil portion 32.

[0012] The orientation of the trip strips 40 and 44 in the cavity 34 also increases heat transfer at the leading edge 30 of the airfoil portion 32. If desired, the trip strips 40 and 44 may be oriented at an angle α of approximately 45 degrees relative to the flow direction 48. The leading edges 54 and 56 of the trip strips 40 and 44 are positioned in the region of highest heat load, in this case the leading edge nose 36. This trip strip orientation permits the creation of a turbulent vortex 49 in the cavity 34. The cooling fluid initially hits the leading edges 54 and 56 of the trip strip and separates from the airfoil surface. The flow then re-attaches downstream of the trip strip leading edges 54 and 56 and moves toward the divider rib 60 between the leading edge cavity 34 and the adjacent cavity 62. As the flow approaches the divider rib 60, it is forced toward the opposite airfoil wall. The flow is directed perpendicular to the pressure side and suction side walls 42 and 46, and meets at the center of the cavity 34. The flow is now forced back towards the leading edge 30 of the airfoil portion 32. The result of this flow migration causes a large vortex 49 that drives flow into the leading edge of the cavity, acting as an impingement jet which also enhances heat transfer at the leading edge nose 36.

[0013] Using the trip strip configuration of the present invention, radial flowing leading edge cavities of turbine engine components will see an increase in convective heat transfer at the leading edge nose of the cavity.

[0014] The particular orientation of the trip strip configuration allows for cooling flow to impinge on the leading edge nose 36, further enhancing heat transfer. The leading edge of the trip strips 40 and 44 is located near the nose 36 of the leading edge cavity 34.

[0015] The trip strips 40, although skewed at an angle α with respect to the direction of flow 48 along the pressure-side wall 42, become normal to the direction of flow 48 as they wrap around the nose 36 of the leading edge cavity 34, increasing the turbulent vortex 49 generated by the trip strips 40 and 44, and subsequently increasing the heat transfer coefficient.

[0016] Away from the leading edge nose 36, the staggered and 45 degree angled trip strips generate a vortex that impinges flow onto the nose 36 of the leading edge cavity.

[0017] The trip strip configuration of the present invention maintains a P/E ratio between 3 and 25 where P is the radial pitch in between trip strips and E is trip strip height. Further, the trip strip configuration described herein maintains an E/H ratio of between 0.15 and 1.50 where E is trip strip height and H is the height of the cavity 34.

[0018] Airflow testing has shown that the heat transfer coefficients at the leading edge of the airfoil adjacent to the external stagnation region when using the staggered trip strips of the present invention are enhanced by approximately two times, greatly increasing airfoil oxidation and thermo-mechanical fatigue cracking life.

Claims

1. A turbine engine component comprising:

an airfoil portion (32) having a leading edge (30), a suction side (46), a pressure side (42) and a trailing edge;

a radial flow leading edge cavity (34) through which a cooling fluid flows for cooling said leading edge and a divider rib (60) between said leading edge cavity (34) and an adjacent cavity (62); and

means for generating a vortex (49) in said leading edge cavity (34) which impinges on a nose portion (36) of said leading edge cavity (34);

wherein said vortex generating means comprises a plurality of first trip strips (40) wrapped around said nose portion (36) of said leading edge cavity (34) and a plurality of second trip strips (44), wherein said first trip strips (40) are mounted to the pressure side (42) of said airfoil portion (32) and said second trip strips (44) are mounted on said suction side (46) of said airfoil portion (32), and wherein said first trip strips (40) are skewed at an angle α with respect to the direction of flow (48) of said cooling fluid along the pressure side (42) of the airfoil portion (32) and become normal to the direction of flow (48) as they wrap around said nose portion (36) of said leading edge cavity (34); wherein

there are a plurality of gaps (47) between adjacent ones of said first trip strips (40; 44) and said second trip strips (44; 40);

characterised in that said plurality of gaps (47) are located along a parting line (70) of said airfoil portion (32).

2. The turbine engine component according to claim 1, wherein said first trip strips (40; 44) and said second trip strips (44; 40) are staggered.

3. The turbine engine component according to claim 1 or 2, wherein said first trip strips (40; 44) and second trip strips (44; 40) are positioned along a direction of flow (48) of said cooling fluid.

4. The turbine engine component according to claim 3, wherein said first and second trip strips (40; 44) are each oriented at an angle of 45 degrees relative to the direction of flow of said cooling fluid.

5. The turbine engine component according to claim 3, wherein each of said first and second trip strips (40; 44) has a leading edge and said leading edge of each of said trip strips is positioned in a region of highest heat load.

6. The turbine engine component according to any of claims 3 to 5, wherein said trip strips (40; 44) have a P/E ratio in the range of from 3 to 25 where P is a radial pitch between trip strips and E is trip strip height.

7. The turbine engine component according to any of claims 3 to 6, wherein said trip strips (40; 44) have an E/H ratio between 0.15 and 1.50 where E is trip strip height and H is height of the cavity.

8. The turbine engine component according to any of claims 3 to 6, wherein said first trip strips (40; 44) and said second trip strips (44; 40) are staggered approximately one half pitch apart.

Ansprüche

1. Turbinenmotorkomponente, umfassend:

einen Schaufelblattabschnitt (32) mit einer Vorderkante (30), einer Ansaugseite (46), einer Druckseite (42) und einer Hinterkante;

einen Radialströmungsvorderkantenhohlraum (34), durch den ein Kühlfluid strömt, um die Vorderkante zu kühlen, und eine Teilungsrippe (60) zwischen dem Vorderkantenhohlraum (34) und einem benachbarten Hohlraum (62); und

ein Mittel zum Erzeugen eines Wirbels (49) in dem Vorderkantenhohlraum (34), der auf einen Nasenabschnitt (36) des Vorderkantenhohlraums (34) trifft;

wobei das Wirbelerzeugungsmittel eine Vielzahl erster Turbulatoren (40), die um den Nasenabschnitt (36) des Vorderkantenhohlraums (34) gewickelt ist, und eine Vielzahl zweiter Turbulatoren (44) umfasst, wobei die ersten Turbulatoren (40) an der Druckseite (42) des Schaufelblattabschnitts (32) angebracht sind und die zweiten Turbulatoren (44) an der Ansaugseite (46) des Schaufelblattabschnitts (32) angebracht sind, und wobei die ersten Turbulatoren (40) mit einem Winkel α in Bezug auf die Strömungsrichtung (48) des Kühlfluids an der Druckseite (42) des Schaufelblattabschnitts (32) abgeschrägt sind und lotrecht zur Strömungsrichtung (48) werden, während sie sich um den Nasenabschnitt (36) des Vorderkantenhohlraums (34) wickeln; wobei

eine Vielzahl von Spalten (47) zwischen benachbarten der ersten Turbulatoren (40; 44) und zweiten Turbulatoren (44; 40) vorliegt;

dadurch gekennzeichnet, dass die Vielzahl von Spalten (47) an einer Trennlinie (70) des Schaufelblattabschnitts (32) angeordnet ist.

2. Turbinenmotorkomponente nach Anspruch 1, wobei die ersten Turbulatoren (40; 44) und zweiten Turbulatoren (44; 40) gestaffelt sind.

3. Turbinenmotorkomponente nach Anspruch 1 oder 2, wobei die ersten Turbulatoren (40; 44) und zweiten Turbulatoren (44; 40) entlang einer Strömungsrichtung (48) des Kühlfluids angeordnet sind.

4. Turbinenmotorkomponente nach Anspruch 3, wobei die ersten Turbulatoren (40; 44) jeweils in einem Winkel von 45 Grad relativ zur Strömungsrichtung des Kühlfluids ausgerichtet sind.

5. Turbinenmotorkomponente nach Anspruch 3, wobei jeder der ersten Turbulatoren (40; 44) eine Vorderkante aufweist und die Vorderkante eines jeden Turbulators in einem Bereich mit der höchsten Wärmelast angeordnet ist.

6. Turbinenmotorkomponente nach einem der Ansprüche 3 bis 5, wobei die Turbulatoren (40; 44) ein P/E-Verhältnis im Bereich von 3 bis 25 aufweisen, wobei P das radiale Abstandsmaß zwischen den Turbulatoren und E die Turbulatorhöhe ist.

7. Turbinenmotorkomponente nach einem der Ansprüche 3 bis 6, wobei die Turbulatoren (40; 44) ein E/H-Verhältnis im Bereich von 0,15 bis 1,50 aufweisen, wobei E die Turbulatorhöhe ist und H die Höhe des Hohlraums ist.

8. Turbinenmotorkomponente nach einem der Ansprüche 3 bis 6, wobei die ersten Turbulatoren (40; 44) und zweiten Turbulatoren (44; 40) um ein halbes Abstandsmaß zueinander gestaffelt sind.

Revendications

1. Composant de moteur à turbine comprenant :

une portion de profil aérodynamique (32) ayant un bord d'attaque (30), un côté aspiration (46), un côté pression (42) et un bord de fuite ;

une cavité de bord d'attaque à écoulement radial (34) à travers laquelle s'écoule un fluide de refroidissement pour refroidir ledit bord d'attaque et une nervure de diviseur (60) entre ladite cavité de bord d'attaque (34) et une cavité adjacente (62) ; et

un moyen pour générer un vortex (49) dans ladite cavité de bord d'attaque (34) qui impacte une portion de nez (36) de ladite cavité de bord d'attaque (34) ;

dans lequel ledit moyen de génération de vortex comprend une pluralité de premières bandes de déclenchement (40) enroulées autour de ladite portion de nez (36) de ladite cavité de bord d'attaque (34) et une pluralité de secondes bandes de déclenchement (44), dans lequel lesdites premières bandes de déclenchement (40) sont montées sur le côté pression (42) de ladite portion de profil aérodynamique (32) et lesdites secondes bandes de déclenchement (44) sont montées sur ledit côté aspiration (46) de ladite portion de profil aérodynamique (32), et dans lequel lesdites premières bandes de déclenchement (40) sont obliques à un angle α par rapport à la direction d'écoulement (48) dudit fluide de refroidissement le long du côté pression (42) de la portion de profil aérodynamique (32) et deviennent normales à la direction d'écoulement (48) lorsqu'elles sont enroulées autour de ladite portion de nez (36) de ladite cavité de bord d'attaque (34) ; dans lequel il y a une pluralité d'écartements (47) entre des bandes adjacentes desdites premières bandes de déclenchement (40 ; 44) et desdites secondes bandes de déclenchement (44 ; 40) ;

caractérisé en ce que ladite pluralité d'écartements (47) est située le long d'une ligne de jonction (70) de ladite portion de profil aérodynamique (32).

2. Composant de moteur à turbine selon la revendication 1, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et lesdites secondes bandes de déclenchement (44 ; 40) sont décalées.

3. Composant de moteur à turbine selon la revendication 1 ou 2, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et secondes bandes de déclenchement (44 ; 40) sont positionnées le long d'une direction d'écoulement (48) dudit fluide de refroidissement.

4. Composant de moteur à turbine selon la revendication 3, dans lequel lesdites premières et secondes bandes de déclenchement (40 ; 44) sont chacune orientées à un angle de 45 degrés par rapport à la direction d'écoulement dudit fluide de refroidissement.

5. Composant de moteur à turbine selon la revendication 3, dans lequel chacune desdites premières et secondes bandes de déclenchement (40 ; 44) a un bord d'attaque et ledit bord d'attaque de chacune desdites bandes de déclenchement est positionné dans une région de charge calorifique la plus élevée.

6. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 5, dans lequel lesdites bandes de déclenchement (40 ; 44) ont un rapport P/E dans la plage de 3 à 25 où P est un pas radial entre des bandes de déclenchement et E est une hauteur de bande de déclenchement.

7. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 6, dans lequel lesdites bandes de déclenchement (40 ; 44) ont un rapport E/H entre 0,15 et 1,50 où E est une hauteur de bande de déclenchement et H est une hauteur de la cavité.

8. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 6, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et lesdites secondes bandes de déclenchement (44 ; 40) sont décalées approximativement d'un demi-pas.

Drawing