Cooling schemes for airfoils for gas turbine engines

1. An airfoil for a gas turbine engine, the airfoil comprising:

a leading edge, a trailing edge, a pressure side exterior wall extending between the leading edge and the trailing edge and defining a pressure side, and a suction side exterior wall extending between the leading edge and the trailing edge and defining a suction side, wherein a plurality of cooling passages are formed within the airfoil;

a plurality of first interior ribs extend from the pressure side exterior wall to the suction side exterior wall; and

a plurality of second interior ribs extend from the suction side exterior wall toward the pressure side exterior wall and intersect with one of the first interior ribs of the plurality of first interior ribs, wherein:

at least one pressure side main body cooling passage is defined between the pressure side exterior wall and two first interior ribs of the plurality of first interior ribs,

at least one suction side main body cooling passage is defined between the suction side exterior wall, a first interior rib, and a second interior rib, and

at least one isolated main body cooling passage is defined between two first interior ribs and two second interior ribs,

wherein the at least one isolated main body cooling passage is configured to supply cooling air to a tip of the airfoil, and

wherein the airfoil defines a camber line extending from the leading edge to the trailing edge, wherein the camber line is defined as a line that is equidistant from the pressure side and the suction side, wherein the camber line passes through each pressure side main body cooling passage.

2. The airfoil of claim 1 , wherein at least one resupply flow aperture is arranged within at least one interior rib such that cooling air is resupplied from the at least one isolated main body cooling passage and at least one of a pressure side main body cooling passage and a suction side main body cooling passage.

3. The airfoil of claim 1 , wherein the plurality of cooling passages further comprises at least one leading edge cooling passage arranged along the leading edge of the airfoil.

4. The airfoil of claim 1 , wherein the plurality of cooling passages further comprises at least one trailing edge cooling passages arranged along the trailing edge of the airfoil.

5. The airfoil of claim 1 , wherein the camber line defines a camber line height H C and the pressure side main body cooling passage has a passage height from a base to an apex defined as 1.1·H C ≤H 1 ≤2.0·H C .

6. The airfoil of claim 1 , wherein the camber line does not intersect any of the suction side main body cooling passages.

7. The airfoil of claim 6 , wherein the camber line defines a camber line height H C and the suction side main body cooling passage has a passage height from a base to an apex defined as 0.1·H C ≤H 1 ≤1.0·H C .

8. The airfoil of claim 1 , further comprising a first heat transfer augmentation feature formed on the pressure side exterior wall within the pressure side main body cooling passage and a second heat transfer augmentation feature formed on the suction side exterior wall within the suction side main body cooling passage.

9. The airfoil of claim 8 , wherein each of the first and second heat transfer augmentation features have a height between 0.005 inches and 0.020 inches.

10. The airfoil of claim 8 , wherein the first heat transfer augmentation feature has a first height E 1 and the pressure side main body cooling passage has a first passage height H 1 , wherein:

E

1

H

1

<

0.05

.

11. The airfoil of claim 8 , wherein the second heat transfer augmentation feature has a second height E 2 and the suction side main body cooling passage has a second passage height H 2 , wherein:

0.50

≥

E

2

H

2

≥

0.10

.

12. The airfoil of claim 8 , wherein each of the first and second heat transfer augmentation features comprises at least one of a normal trip strip, a skewed trip strip, a chevron trip strip, a hemispherical protrusion, and a pin fin.

13. The airfoil of claim 1 , wherein a cross-sectional area of the at least one pressure side main body cooling passage in a flow direction through the at least one pressure side main body cooling passage is greater than a cross-sectional area of the at least one suction side main body cooling passage in a flow direction through the at least one suction side main body cooling passage.

14. A gas turbine engine comprising:

a turbine section;

a compressor section; and

an airfoil located in at least one of the turbine section and the compressor section, the airfoil comprising:

a leading edge, a trailing edge, a pressure side exterior wall extending between the leading edge and the trailing edge and defining a pressure side, and a suction side exterior wall extending between the leading edge and the trailing edge and defining a suction side, wherein a plurality of cooling passages are formed within the airfoil;

a plurality of first interior ribs extend from the pressure side exterior wall to the suction side exterior wall; and

a plurality of second interior ribs extend from the suction side exterior wall toward the pressure side exterior wall and intersect with a first interior rib, wherein:

at least one pressure side main body cooling passage is defined between the pressure side exterior wall and two first interior ribs of the plurality of first interior ribs,

at least one suction side main body cooling passage is defined between the suction side exterior wall, a first interior rib, and a second interior rib, and

at least one isolated main body cooling passage is defined between two first interior ribs and two second interior ribs,

wherein the at least one isolated main body cooling passage is configured to supply cooling air to a tip of the airfoil, and

wherein the airfoil defines a camber line extending from the leading edge to the trailing edge, wherein the camber line is defined as a line that is equidistant from the pressure side and the suction side, wherein the camber line passes through each pressure side main body cooling passage.

15. The gas turbine engine of claim 14 , wherein at least one resupply flow aperture is arranged within at least one interior rib such that cooling air is resupplied from the at least one isolated main body cooling passage and at least one of a pressure side main body cooling passage and a suction side main body cooling passage.

16. The gas turbine engine of claim 14 , wherein the plurality of cooling passages further comprises at least one of a leading edge cooling passage arranged along the leading edge of the airfoil and a trailing edge cooling passage arranged along the trailing edge of the airfoil.

17. The gas turbine engine of claim 14 , wherein the camber line does not intersect any of the suction side main body cooling passages.

18. The gas turbine engine of claim 14 , further comprising a first heat transfer augmentation feature formed on the pressure side exterior wall within the pressure side main body cooling passage and a second heat transfer augmentation feature formed on the suction side exterior wall within the suction side main body cooling passage.

19. The gas turbine engine of claim 14 , wherein a cross-sectional area of the at least one pressure side main body cooling passage in a flow direction through the at least one pressure side main body cooling passage is greater than a cross-sectional area of the at least one suction side main body cooling passage in a flow direction through the at least one suction side main body cooling passage.

20. The airfoil of claim 1 , wherein the supply of cooling air to the tip of the airfoil from the isolated main body cooling passage is through one of a purge hole in tip of the airfoil or into a tip flag cooling passages of the airfoil.