Double wall turbine gas turbine engine blade cooling configuration
An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a tip. A skin 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 core passage and the skin passage are configured to have a same direction of fluid flow. A resupply hole fluidly interconnects the core and skin passages. A centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage. The resupply hole has an exit at the skin passage and the exit has a diffuser.
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 core cooling passage is arranged between the pressure and suction side walls in a thickness direction and extends radially toward a tip, a skin 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 core passage and the skin passage are configured to have a direction of fluid flow that is the same as one another, and a resupply hole fluidly interconnects the core and skin passages, a centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage, wherein the resupply hole has a metering section extending a length along the centerline of the resupply hole, the metering section is upstream from and leads into an exit at the skin passage, the exit has a diffuser, wherein the skin passage has an aspect ratio in a range of 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).
2. The airfoil of claim 1 , wherein the acute angle is in a range of 5°-45°.
3. The airfoil of claim 1 , wherein the diffuser provides an aperture provided on an inner surface of the cold side wall facing the hot side wall, the aperture includes an aperture length and an aperture width, and a ratio of the aperture length to the aperture width in a range of 0.5:1 to 3:1.
4. The airfoil of claim 3 , wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections providing rounded corners.
5. The airfoil of claim 3 , wherein the diffuser includes multiple lobes having peaks and valleys extending in a direction of the centerline of the resupply hole, the peaks and valleys providing scallops in the inner surface.
6. The airfoil of claim 3 , wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections, the bottom wall narrowing to a neck.
7. The airfoil of claim 1 , comprising a serpentine cooling passage having first, second and third cooling passages, the first and third cooling passages having a direction of fluid flow toward the tip, and the second cooling passage having a direction of fluid flow away from the tip, the core passage provided by one of the first, second and third cooling passages.
8. The airfoil of claim 1 , wherein a film cooling hole extends from the skin passage to the exterior airfoil surface.
9. The airfoil of claim 1 , wherein the airfoil is a turbine blade.
10. A gas turbine engine comprising:
a combustor section arranged fluidly between compressor and turbine sections; and
an airfoil arranged in the turbine section, the airfoil including 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 core cooling passage is arranged between the pressure and suction side walls in a thickness direction and extends radially toward a tip, a skin 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 core passage and the skin passage are configured to receive a cooling fluid from the compressor section and have a direction of fluid flow that is the same as one another, and a resupply hole fluidly interconnects the core and skin passages, a centerline of the resupply hole is arranged at an acute angle relative to the direction of fluid flow in the core passage and is configured to provide a low turbulence flow region in the skin passage, wherein the resupply hole has a metering section extending a length along the centerline of the resupply hole, the metering section is upstream from and leads into an exit at the skin passage, the exit has a diffuser, wherein the skin passage has an aspect ratio in a range of 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).
11. The gas turbine engine of claim 10 , wherein the acute angle is in a range of 5°-45°.
12. The gas turbine engine of claim 10 , wherein the diffuser provides an aperture in an inner surface provided on an inner surface of the cold side wall facing the hot side wall, the aperture includes an aperture length and an aperture width, and a ratio of the aperture length to the aperture width in a range of 0.5:1 to 3:1.
13. The gas turbine engine of claim 12 , wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections providing rounded corners.
14. The gas turbine engine of claim 12 , wherein the diffuser includes multiple lobes having peaks and valleys extending in a direction of the centerline of the resupply hole, the peaks and valleys providing scallops in the inner surface.
15. The gas turbine engine of claim 12 , wherein the diffuser has a bottom wall joined to laterally spaced apart side walls at intersections, the bottom wall narrowing to a neck.
16. The gas turbine engine of claim 10 , wherein the airfoil is a turbine blade, comprising a serpentine cooling passage having first, second and third cooling passages, the first and third cooling passages having a direction of fluid flow toward the tip, and the second cooling passage having a direction of fluid flow away from the tip, the core passage provided by one of the first, second and third cooling passages.
17. The gas turbine engine of claim 10 , wherein a film cooling hole extends from the skin passage to the exterior airfoil surface.