Double wall turbine gas turbine engine vane platform cooling configuration with main core resupply

1. An airfoil comprising:

pressure and suction side walls extending in a chord-wise direction between leading and trailing edges, the pressure and suction side walls extending in a radial direction to provide an exterior airfoil surface, a main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform, an airfoil skin cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall, the hot side wall defines a portion of the exterior airfoil surface, and the cold side wall defines a portion of the core passage, the airfoil skin cooling passage extends to a platform skin cooling passage arranged in the platform, the airfoil skin cooling passage is configured to receive a cooling fluid from a location upstream from the platform skin cooling passage, the platform skin cooling passage terminates at a fluid exit provided at the perimeter of the platform, the perimeter includes at least one of axial and circumferential mate faces, a resupply hole that is configured to supply from the main-body core cooling passage to at least one of the airfoil skin cooling passage and the platform skin cooling passage, wherein the airfoil skin cooling passage has an aspect ratio that may vary between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm, and the platform skin cooling passage has an aspect ratio that may vary between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm), wherein the resupply hole is arranged at a bend fluidly connecting the airfoil skin cooling passage and the platform skin cooling passage, the resupply hole is configured to supplement the cooling fluid with a resupply air from the main-body core cooling passage.

2. The airfoil of claim 1 , wherein the fluid exit is provided on at least one of the axial and circumferential mate faces.

3. The airfoil of claim 1 , wherein at least one of the airfoil skin cooling passage and the platform skin cooling passage includes trip strips, turbulators, pin fins and/or dimples.

4. The airfoil of claim 1 , wherein the airfoil is a turbine vane.

5. The airfoil of claim 4 , wherein the resupply hole has a ratio of length to hydraulic diameter (L/Dh) in a range of 1-20.

6. The airfoil of claim 1 , wherein a film cooling hole extends from the airfoil skin cooling passage to the exterior airfoil surface.

7. A gas turbine engine comprising:

a combustor section arranged fluidly between a compressor section and a turbine section;

a stationary stage and a rotating stage arranged in one of the compressor and turbine sections;

an airfoil arranged in one of the stationary and rotating stages, the airfoil includes pressure and suction side walls extending in a chord-wise direction between leading and trailing edges, the pressure and suction side walls extending in a radial direction to provide an exterior airfoil surface, a main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform, an airfoil skin cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall, the hot side wall defines a portion of the exterior airfoil surface, and the cold side wall defines a portion of the core passage, the airfoil skin cooling passage extends to a platform skin cooling passage arranged in the platform, a resupply hole that is configured to supply fluid from the main-body core cooling passage to at least one of the airfoil skin cooling passage and the platform skin cooling passage, wherein the airfoil skin cooling passage has an aspect ratio that may vary between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm, and the platform skin cooling passage has an aspect ratio that may vary between 3:1≥H/W≥1:5, wherein H corresponds to a passage height and W corresponds to a passage width, wherein the passage height (H) is in a range of 0.010-0.200 inches (0.25-5.08 mm), wherein the resupply hole is arranged at a bend fluidly connecting the airfoil skin cooling passage and the platform skin cooling passage; and

a cooling source configured to provide a cooling fluid, wherein the cooling fluid is supplied to the airfoil skin cooling passage at a location upstream from the platform skin cooling passage, the platform skin cooling passage terminates at a fluid exit provided at the perimeter of the platform, the perimeter includes at least one of axial and circumferential mate faces, the resupply hole is configured to supplement the cooling fluid with a resupply air from the main-body core cooling passage.

8. The gas turbine engine of claim 7 , wherein the fluid exit is provided on at least one of the axial mate face and the circumferential mate face.

9. The gas turbine engine of claim 7 , wherein the fluid exit is provided on the axial mate face or the circumferential mate face.

10. The gas turbine engine of claim 7 , wherein the airfoil is a turbine vane arranged in the stationary stage of the turbine section.

11. The gas turbine engine of claim 10 , wherein the resupply hole has a ratio of length to hydraulic diameter (L/Dh) in a range of 1-20.

12. The gas turbine engine of claim 7 , wherein a film cooling hole extends from the airfoil skin cooling passage to the exterior airfoil surface.