Turbine airfoil film cooling holes
A turbomachinery component includes an airfoil section with a cooling passage extending within the airfoil section. A wall is defined between the cooling passage and an exterior surface of the airfoil. At least one row of cooling holes is positioned along a cooling portion of the external wall proximate a leading edge of the airfoil for fluid communication between the cooling passage and the exterior surface of the airfoil. The cooling portion of the airfoil wall is thicker than an average adjacent airfoil wall thickness.
1. A turbomachinery component, comprising:
an airfoil section;
a cooling passage extending within the airfoil section;
a wall defined between the cooling passage and an exterior surface of the airfoil section; and
at least one row of cooling holes along a cooling portion of the wall proximate a leading edge of the airfoil section for fluid communication between the cooling passage and the exterior surface of the airfoil section, wherein an internal surface of the wall adjacent to the cooling passage has a convex shape and a thickness of the convex shape is greater than portions of the wall on each side of the convex shape.
2. The turbomachinery component of claim 1 , wherein the thickness of the convex shape is at least 50% thicker than an average thickness of the portions of the wall on each side of the convex shape and at most 400% greater than the average thickness of the portions of the wall on each side of the convex shape.
3. The turbomachinery component of claim 1 , wherein the portions of the wall on each side of the convex shape include walls of a suction side and a pressure side of the airfoil section.
4. The turbomachinery component of claim 1 , wherein a majority of the at least one row of cooling holes have a L/D ratio range of 5-20.
5. The turbomachinery component of claim 1 , wherein the at least one row of cooling holes is rows of cooling holes radially spaced along the cooling portion.
6. The turbomachinery component of claim 1 , including one row of cooling holes.
7. The turbomachinery component of claim 1 , including two rows of cooling holes.
8. The turbomachinery component of claim 1 , including three rows of cooling holes.
9. The turbomachinery component of claim 1 , including four rows of cooling holes.
10. The turbomachinery component of claim 1 , including five rows of cooling holes.
11. The turbomachinery component of claim 1 , wherein the number of cooling holes of a first cooling hole row at a first airfoil radial location is different from a number of cooling holes at a second cooling hole row at a second airfoil radial location.
12. The turbomachinery component of claim 1 , wherein the at least one row of cooling holes is a plurality of rows and one of the plurality of rows extends radially further than another one of the plurality of rows across the airfoil section.
13. The turbomachinery component of claim 1 , further comprising a row of gill holes along a suction side of the airfoil section relative to the at least one row of cooling holes.
14. The turbomachinery component of claim 1 , further comprising a row of gill holes along a pressure side of the airfoil section relative to the at least one row of cooling holes.
15. The turbomachinery component of claim 1 , wherein the airfoil section is a turbine blade.
16. The turbomachinery component of claim 1 , wherein the airfoil section is a turbine vane.
17. The turbomachinery component of claim 1 , where the at least one row of cooling holes are positioned at the leading edge of the airfoil section.
18. The turbomachinery component of claim 1 , wherein the at least one row of cooling holes include holes with a constant cross-sectional area from an inlet of the cooling passage to an exit of the exterior surface of the airfoil section.
19. The turbomachinery component of claim 1 , wherein the at least one row of cooling holes include holes with a round metering section and a diffusing section.