Patent ID: 6916155

Claim:
A gas turbine blade for an airplane engine, the blade comprising in its central portion first and second central cooling circuits, each central cooling circuits comprising at least first and second cavities extending radially on a concave side of the blade, at least one cavity extending on a convex side of the blade, an air admission opening at a radial end of the first concave side cavity for feeding the central cooling circuit with cooling air, a first passage putting the other radial end of the first concave side cavity into communication with an adjacent radial end of the convex side cavity, a second passage putting the other radial end of the convex side cavity into communication with an adjacent radial end of the second concave side cavity, and outlet orifices opening out into the second concave side cavity and through a concave face of the blade, wherein the first and second central cooling circuits are substantially symmetrical to each other about a plane transverse to the blade.