Directional cooling arrangement for airfoils

1. An airfoil, comprising:

an airfoil section having an internal wall and an external wall, the external wall defining pressure and suction sides between a leading edge and a trailing edge, and the airfoil section defining a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section;

a first cavity and a second cavity separated by the internal wall, the second cavity bounded by the external wall at the leading edge;

at least one crossover passage within the internal wall that connects the first cavity to the second cavity; and

wherein the at least one crossover passage defines a passage axis, the passage axis defining a passage angle with respect to the mean camber line such that the passage axis extends transversely from the mean camber line to intersect a surface of the second cavity.

2. The airfoil as recited in claim 1 , wherein the passage angle is between 45 degrees and 75 degrees.

3. The airfoil as recited in claim 2 , wherein the passage axis intersects the suction side.

4. The airfoil as recited in claim 2 , wherein the passage axis intersects the pressure side.

5. The airfoil as recited in claim 1 , wherein the airfoil section extends in a spanwise direction from a 0% span position to a 100% span position, and an outlet of the at least one crossover passage is defined at a span position that is between an 80% span position and the 100% span position.

6. The airfoil as recited in claim 5 , wherein the passage angle is between 15 degrees and 90 degrees.

7. The airfoil as recited in claim 1 , comprising a plurality of film cooling passages within the external wall and extending from a surface of the second cavity, wherein the passage axis intersects the external wall at an intersection point, and the second cavity is free of any film cooling passages aft of the intersection point.

8. The airfoil as recited in claim 7 , wherein the mean camber line intersects the at least one crossover passage.

9. The airfoil as recited in claim 8 , wherein the passage axis intersects the suction side.

10. The airfoil as recited in claim 7 , wherein the at least one crossover passage extends between an inlet and an outlet, the outlet spaced apart from the external wall.

11. The airfoil as recited in claim 1 , wherein the airfoil is a turbine blade.

12. A casting core for an airfoil, comprising:

a first portion corresponding to a leading edge cavity of an airfoil and a second portion corresponding to a second cavity of the airfoil, the airfoil defining a mean camber line that bisects a thickness of the airfoil between pressure and suction sides of the airfoil;

a plurality of connectors coupling the first portion and the second portion, the plurality of connectors corresponding to a plurality of crossover passages of the airfoil, each of the plurality of connectors defining a passage axis; and

wherein the passage axis of each of the plurality of connectors is arranged at a corresponding passage angle relative to the mean camber line, the passage angle being between 15 degrees and 90 degrees.

13. The casting core for an airfoil as recited in claim 12 , wherein the mean camber line intersects at least one of the plurality of crossover passages.

14. The casting core for an airfoil as recited in claim 12 , wherein the airfoil is a turbine airfoil.

15. A gas turbine engine, comprising:

an array of airfoils circumferentially distributed about an engine axis, each airfoil of the array of airfoils including an airfoil section having pressure and suction sides between a leading edge and a trailing edge, the airfoil section defining a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section, facing pressure and suction sides of adjacent airfoils defining a throat, the throat defined as a minimum distance between the facing pressure and suction sides at a respective span position; and

wherein airfoil section comprises:

an external wall and an internal wall;

a first cavity and a second cavity separated by the internal wall, the second cavity bounded by the external wall at the leading edge;

a plurality of crossover passages within the internal wall and connecting the first cavity to the second cavity, each of the plurality of crossover passages defining a passage axis; and

wherein the passage axis of each of the plurality of crossover passages is arranged at a corresponding passage angle relative to the mean camber line such that the passage axis of each of the plurality of crossover passages intersects the external wall at a location between the throat and an aftmost location of the second cavity with respect to the engine axis.

16. The gas turbine engine as recited in claim 15 , wherein the airfoil section extends in a radial direction with respect to the engine axis from a 0% span position to a 100% span position, and an outlet of each of the plurality of crossover passage is defined at a respective span position that is between an 80% span position and the 100% span position.

17. The gas turbine engine as recited in claim 16 , comprising a plurality of film cooling passages within the external wall and extending from a surface of the second cavity, wherein the passage axis of at least some of the plurality of crossover passages intersects the external wall at an intersection point, and the second cavity is free of any film cooling passages aft of the intersection point.

18. The gas turbine engine as recited in claim 17 , wherein the passage axis of each of the plurality of crossover passages intersects the suction side.

19. The gas turbine engine as recited in claim 15 , comprising a compressor section and a turbine section, with the array of airfoils located in at least one of the compressor section and the turbine section.

20. The gas turbine engine as recited in claim 19 , wherein the array of airfoils are rotatable blades.