Double wall turbine gas turbine engine vane platform cooling configuration with baffle impingement

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 is configured to receive a cooling fluid and extends to a platform skin cooling passage arranged in the platform, the platform skin cooling passage terminates at a fluid exit provided at the perimeter of the platform, the perimeter includes axial and circumferential mate faces the platform skin cooling passage resupplied by a resupply hole arranged at a backside of the platform, the resupply hole in communication with the platform skin cooling passage, wherein the airfoil skin cooling passage feeds the cooling fluid to the platform cooling passage, and the resupply hole is arranged fluidly between the fluid exit and the airfoil skin cooling passage to supplement the cooling fluid with a resupply air from the backside of the platform.

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 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).

5. The airfoil of claim 1 , wherein 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).

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

7. The airfoil of claim 6 , wherein the platform includes a rail, and a baffle plate is supported by the rail and provides a cavity.

8. The airfoil of claim 7 , wherein the platform is one of an inner platform and an outer platform, the baffle plate includes holes in fluid communication with the platform cooling passage, the holes in fluid communication with the cavity.

9. The airfoil of claim 7 , wherein the backside includes a channel, and no gap is provided between the platform wall and the baffle plate in an area immediately adjacent to the channel.

10. The airfoil of claim 1 , wherein the airfoil skin cooling passage and the platform skin cooling passage are 90°+/−20° relative to one another.

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

12. 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 providing a vane arranged in the stationary stage, 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, the platform skin cooling passage resupplied by a resupply hole arranged at a backside of the platform;

the platform skin cooling passage terminates at a fluid exit provided at the perimeter of the platform, the perimeter includes axial and circumferential mate faces, and the fluid exit is provided on at least one of one of the axial and circumferential mate faces;

both of the platform skin cooling passage and the platform skin cooling passage have 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);

the airfoil skin cooling passage and the platform skin cooling passage are 90°+/−20° relative to one another; and

a cooling source configured to provide a cooling fluid, wherein the cooling fluid is supplied to the airfoil skin cooling passage, and the airfoil skin cooling passage feeds the cooling fluid to the platform cooling passage, the resupply hole is arranged fluidly between the fluid exit and the airfoil skin cooling passage to supplement the cooling fluid with a resupply air from the backside of the platform.

13. The gas turbine engine of claim 12 , wherein the airfoil is a turbine vane arranged in the turbine section.

14. The gas turbine engine of claim 13 , wherein the platform is one of an inner platform and an outer platform, a baffle plate supported by the platform includes holes in fluid communication with a cavity provided between the baffle plate and the platform.

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