Airfoil having improved throughflow cooling scheme and damage resistance
Vanes for gas turbine engines are described. The vanes include an airfoil body extending between a first platform and a second platform, the airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side and at least one skin core cavity comprising a first skin core cavity, wherein the first skin core cavity defines a throughflow passage through the vane between the first platform and the second platform, wherein the at least one skin core cavity is defined between an external hot wall of the vane and an internal cold wall of the vane.
1. A vane for a gas turbine engine, the vane comprising:
an airfoil body extending between a first platform and a second platform, the airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side;
a first main body cavity arranged along a portion of an external hot wall of the pressure side of the airfoil body, the first main body cavity defining a first main body cavity internal cold wall;
a plurality of suction side skin core cavities defined within the airfoil body along the suction side thereof, wherein at least one suction side skin core cavity of the plurality of suction side skin core cavities defines a throughflow passage through the airfoil body between the first platform and the second platform, wherein each suction side skin core cavity is defined between an external hot wall of the suction side and the first main body cavity internal cold wall;
a second main body cavity aft of the first main body cavity and arranged along a portion of the external hot wall of the suction side of the airfoil body, the second main body cavity defining a second main body cavity internal cold wall;
a plurality of pressure side skin core cavities defined within the airfoil body along the pressure side thereof, wherein at least one pressure side skin core cavity of the plurality of pressure side skin core cavities defines a throughflow passage through the airfoil body between the first platform and the second platform, wherein each pressure side skin core cavity is defined between the external hot wall of the pressure side and the second main body cavity internal cold wall; and
a third main body cavity located aft of the second main body cavity, wherein the first main body cavity, the second main body cavity, and the third main body cavity define a serpentine flow path through the airfoil body.
2. The vane of claim 1 , wherein each cavity of the plurality of suction side skin core cavities is a throughflow passage between the first platform and the second platform.
3. The vane of claim 1 , wherein each cavity of the plurality of pressure side skin core cavities is a throughflow passage between the first platform and the second platform.
4. The vane of claim 1 , wherein the plurality of suction side skin core cavities are located forward of the second main body cavity along the suction side of the airfoil body.
5. The vane of claim 1 , wherein a flow direction through the each of the suction side skin core cavities and each of the pressure side skin core cavities are the same.
6. The vane of claim 1 , wherein one of the plurality of suction side skin core cavities is a non-throughflow cavity that is connected to an external surface of the vane by at least one exit cooling hole.
7. The vane of claim 1 , wherein one of the plurality of pressure side skin core cavities is a non-throughflow cavity that is connected to an external surface of the vane by at least one exit cooling hole.
8. The vane of claim 1 , wherein the plurality of pressure side skin core cavities are located aft of the first main body cavity along the pressure side of the airfoil body.
9. The vane of claim 1 , wherein the suction side skin core cavities are smooth-walled cavities.
10. The vane of claim 1 , wherein the pressure side skin core cavities are smooth-walled cavities.
11. The vane of claim 1 , wherein at least one of the suction side skin core cavities or the pressure side skin core cavities include at least one heat transfer augmentation feature.
12. A gas turbine engine comprising:
a turbine section having a plurality of airfoils, wherein at least one airfoil comprises a vane having:
an airfoil body extending between a first platform and a second platform, the airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side;
a first main body cavity arranged along a portion of an external hot wall of the pressure side of the airfoil body, the first main body cavity defining a first main body cavity internal cold wall;
a plurality of suction side skin core cavities defined within the airfoil body along the suction side thereof, wherein at least one suction side skin core cavity of the plurality of suction side skin core cavities defines a throughflow passage through the airfoil body between the first platform and the second platform, wherein each suction side skin core cavity is defined between an external hot wall of the suction side and the first main body cavity internal cold wall;
a second main body cavity aft of the first main body cavity and arranged along a portion of the external hot wall of the suction side of the airfoil body, the second main body cavity defining a second main body cavity internal cold wall;
a plurality of pressure side skin core cavities defined within the airfoil body along the pressure side thereof, wherein at least one pressure side skin core cavity of the plurality of pressure side skin core cavities defines a throughflow passage through the airfoil body between the first platform and the second platform, wherein each pressure side skin core cavity is defined between the external hot wall of the pressure side and the second main body cavity internal cold wall; and
a third main body cavity located aft of the second main body cavity, wherein the first main body cavity, the second main body cavity, and the third main body cavity define a serpentine flow path through the airfoil body.
13. The gas turbine engine of claim 12 , wherein each cavity of the plurality of suction side skin core cavities is a throughflow passage between the first platform and the second platform.
14. The gas turbine engine of claim 12 , wherein each cavity of the plurality of pressure side skin core cavities is a throughflow passage between the first platform and the second platform.
15. The gas turbine engine of claim 14 , wherein a flow direction through the each of the suction side skin core cavities and each of the pressure side skin core cavities are the same.
16. The gas turbine engine of claim 12 , wherein one of the plurality of suction side skin core cavities is a non-throughflow cavity that is connected to an external surface of the vane by at least one exit cooling hole.
17. The gas turbine engine of claim 12 , wherein one of the plurality of pressure side skin core cavities is a non-throughflow cavity that is connected to an external surface of the vane by at least one exit cooling hole.
18. The gas turbine engine of claim 12 , wherein the suction side skin core cavities are smooth-walled cavities.
19. The gas turbine engine of claim 12 , wherein the pressure side skin core cavities are smooth-walled cavities.
20. The gas turbine engine of claim 12 , wherein at least one of the suction side skin core cavities or the pressure side skin core cavities include at least one heat transfer augmentation feature.