Cooled blades for a gas turbine engine

Abstract From the Canadian patent

An internally cooled turbine blade for a gas turbine engine is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length. Trip strips may be employed to augment the pressure drop distribution