Gas turbine engine airfoil cooling passage configuration

A gas turbine engine comprising:

• compressor and turbine sections;
• and
• an airfoil in one of the compressor and turbine sections, the airfoil extending in a radial direction from a support, the airfoil body having pressure and suction side walls joined at leading and trailing edges to provide an exterior airfoil surface, a chord-wise direction extending between the leading and trailing edges, and a thickness direction transverse to chord-wise direction and extending between the pressure and suction side walls, cooling passages extending from the support into the airfoil body, the cooling passages including adjacent passageways in the thickness direction and separated by a chord-wise wall, one of the adjacent passageways adjacent to another passageway in the chord-wise direction and separated by a rib in the thickness direction, the rib is discontinued at a location along the radial direction but leaving a radially extending rib portion in the thickness direction near either the pressure side wall or the suction side wall to provide an opening that fluidly connects the one of the adjacent passageways to the other passageway such that the other passageway is wider than the one of the adjacent passageways in the thickness direction, wherein the adjacent passageways are configured to carry fluid from the root in the same direction.