Double wall turbine gas turbine engine blade cooling configuration

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 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 same direction of fluid flow, and a resupply hole fluidly interconnects the core and skin passages, the resupply hole has a single inlet that is fluidly connected to multiple discrete outlet apertures, a centerline of the single inlet is arranged at an 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 angle is in a range of 5° -45°, wherein the skin 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).