Impingement cooling arrangement for airfoils

1. An airfoil for a gas turbine engine comprising:

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

a first impingement cavity and a second impingement cavity, both the first and second impingement cavities bounded by the external wall at a leading edge region that defines the leading edge, a first line perpendicular to the mean camber line intersecting both the first and second impingement cavities, and the leading edge region being a region of the airfoil section from the leading edge and within 20% of a distance along the external wall between the leading and trailing edges;

a first feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall;

a first crossover passage within the internal wall that connects the first impingement cavity and the first feeding cavity, wherein the first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity; and

a second crossover passage within the internal wall connected to the second impingement cavity, wherein the second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

2. The airfoil as recited in claim 1 , wherein the airfoil section extending in a spanwise direction from a 0% span position to a 100% span position, the airfoil is a high lift airfoil, a camber angle is defined by a tangential projection of the mean camber line at the leading and trailing edges, and the camber angle is less than 15 degrees between a 70% span position and the 100% span position.

3. The airfoil as recited in claim 2 , wherein the airfoil is a turbine blade.

4. The airfoil as recited in claim 2 , wherein a rib separating the first impingement cavity and the second impingement cavity is transverse to the mean camber line.

5. The airfoil as recited in claim 2 , wherein a length of a rib separating the first impingement cavity and the second impingement cavity follows the mean camber line.

6. The airfoil as recited in claim 5 , wherein the second crossover passage connects the second impingement cavity and a second feeding cavity, the first and second feeding cavity separated by the length of the rib.

7. An airfoil for a gas turbine engine comprising:

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

a first impingement cavity and a second impingement cavity, both the first and second impingement cavities bounded by the external wall at a leading edge region that defines the leading edge, a first line perpendicular to the mean camber line intersecting both the first and second impingement cavities, and the leading edge region being a region of the airfoil section from the leading edge and within 20% of a distance along the external wall between the leading and trailing edges;

a first feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall;

a first crossover passage within the internal wall that connects the first impingement cavity and the first feeding cavity, wherein the first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity;

a second crossover passage within the internal wall connected to the second impingement cavity, wherein the second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity; and

wherein the first passage axis intersects the external wall at an intersection point, the intersection point being between the trailing edge and a gauge point defined by the airfoil.

8. The airfoil as recited in claim 7 , wherein the second crossover passage connects the second impingement cavity and a second feeding cavity, the second feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall.

9. The airfoil as recited in claim 7 , wherein the first passage axis intersects the surface of the first impingement cavity adjacent to the pressure side, and the second passage axis intersects the surface of the second impingement cavity adjacent to the suction side.

10. The airfoil as recited in claim 7 , wherein:

the first crossover passage extends between a first inlet and a first outlet, the second crossover passage extends between a second inlet and a second outlet, and the first inlet is forward of the second outlet with respect to the chordwise direction; and

the first outlet is forward of the second inlet with respect to the chordwise direction.

11. The airfoil as recited in claim 10 , wherein the second crossover passage connects the second impingement cavity and the first feeding cavity.

12. The airfoil as recited in claim 7 , wherein the first impingement cavity defines a first volume and the second impingement cavity defines a second volume that is less than the first volume.

13. 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 an internal wall and an external wall, the external wall defining pressure and suction sides extending in a chordwise direction between a leading edge and a trailing edge; and

wherein the airfoil section comprises:

a first impingement cavity and a second impingement cavity, both the first and second impingement cavities bounded by the external wall at a leading edge region that defines the leading edge;

a first feeding cavity and a second feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall;

a first set of crossover passages within the internal wall that connect the first impingement cavity and the first feeding cavity, wherein each passage of the first set of crossover passages defines a first passage axis that intersects a surface of the first impingement cavity; and

a second set of crossover passages within the internal wall that connect the second impingement cavity and the second feeding cavity, wherein each passage of the second set of crossover passages defines a second passage axis that intersects a surface of the second impingement cavity;

wherein facing pressure and suction sides of adjacent airfoils define a throat that corresponds to a gauge point, the throat defined as a minimum distance between the facing pressure and suction sides at a respective span position; and

wherein the first passage axis of at least some passages of the first set of crossover passages intersects the external wall at an intersection point, the intersection point being within 10% of a distance from the gauge point with respect to a length between the gauge point and one of the leading edge and the trailing edge.

14. The gas turbine engine as recited in claim 13 , comprising a compressor section and a turbine section, wherein the array of airfoils are located in at least one of the compressor section and the turbine section.

15. The gas turbine engine as recited in claim 13 , wherein the airfoil section defines a mean camber line extending between the leading and trailing edges to bisect a thickness of the airfoil section, both the first and second impingement cavities are bounded by the external wall at the leading edge region, a first line perpendicular to the mean camber line intersecting both the first and second impingement cavities and the leading edge region being a region of the airfoil section from the leading edge and within 20% of a distance along the external wall between the leading and trailing edges.

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

17. The gas turbine engine as recited in claim 16 , wherein the array of airfoils are rotatable blades in the turbine section.

18. 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 an internal wall and an external wall, the external wall defining pressure and suction sides extending in a chordwise direction 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; and

wherein the airfoil section comprises:

a first impingement cavity and a second impingement cavity both bounded by the external wall at a leading edge region that defines the leading edge, a first line perpendicular to the mean camber line intersecting both the first and second impingement cavities, and the leading edge region being a region of the airfoil section from the leading edge and within 20% of a distance along the external wall between the leading and trailing edges;

a first set of crossover passages within the internal wall that connect the first impingement cavity and a first feeding cavity, wherein each passage of the first set of crossover passages defines a first passage axis that intersects a surface of the first impingement cavity; and

a second set of crossover passages within the internal wall, wherein each passage of the second set of crossover passages defines a second passage axis that intersects a surface of the second impingement cavity.

19. The gas turbine engine as recited in claim 18 , wherein:

facing pressure and suction sides of adjacent airfoils define a throat that corresponds to a gauge point, the throat defined as a minimum distance between the facing pressure and suction sides at a respective span position; and

the first passage axis of at least some passages of the first set of crossover passages intersects the external wall at an intersection point, the intersection point being within 10% of a distance from the gauge point with respect to a length between the gauge point and one of the leading edge and the trailing edge.

20. The gas turbine engine as recited in claim 19 , wherein:

the airfoil section includes a second feeding cavity, the first feeding cavity and the second feeding cavity separated from the first impingement cavity and from the second impingement cavity by the internal wall; and

the second set of crossover passages connect the second impingement cavity and the second feeding cavity.

21. The gas turbine engine as recited in claim 19 , wherein:

each passage of the first set of crossover passages extends between a first inlet and a first outlet, each passage of the second set of crossover passages extends between a second inlet and a second outlet, and the first inlet is forward of the second outlet with respect to the chordwise direction;

the second set of crossover passages connect the second impingement cavity and the first feeding cavity, and the first outlet is forward of the second inlet with respect to the chordwise direction.