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 at an outer surface and a cold side wall at an inner surface, wherein the skin passage extends a height in the radial direction, a width in a width direction, and a thickness in a thickness direction, wherein the thickness is less than the width, the hot side wall defines a portion of the exterior airfoil surface, and the cold side wall defines a portion of the core passage at a core passage surface, the core passage and the skin passage are configured to have a same direction of fluid flow, and a resupply hole fluidly interconnects the core and skin passages, a centerline of the resupply hole is arranged at a first angle relative to the direction of fluid flow in the core passage, the first angle lies in a plane parallel to the thickness direction, the centerline of the resupply hole is arranged at a second angle relative to the direction of fluid flow in the core passage, the second angle lies in a plane parallel to the width direction, the second angle is at an acute angle, and the first and second angles are configured to provide a low turbulence flow region in the skin passage.