Abstract: A gas turbine blade of a turbomachine includes in its central portion a centrally-located first cooling circuit at least a suction side cavity, at least a pressure side cavity, at least a central cavity extending between the suction side cavity and the pressure side cavity, a first air admission opening at a radially bottom end of the suction side cavity, a second air admission opening at a radially bottom end of the pressure side cavity, at least a first passage putting a radially top end of the suction side cavity into communication with a radially top end of the central cavity, at least a second passage putting a radially top end of the pressure side cavity into communication with the radially top end of the central cavity, and outlet orifices opening out both into the central cavity and into the pressure side face of the blade.

Abstract: A gas turbine blade for an airplane engine, comprising in its central portion at least a first central cooling circuit comprising at least first and second cavities extending radially on the concave side of the blade, at least one cavity extending on the convex side of the blade, an air admission opening at a radial end of the first concave side cavity for feeding the first 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 the concave face of the blade.

Abstract: A gas turbine blade of a turbomachine includes in its central portion a centrally-located first cooling circuit at least a suction side cavity, at least a pressure side cavity, at least a central cavity extending between the suction side cavity and the pressure side cavity, a first air admission opening at a radially bottom end of the suction side cavity, a second air admission opening at a radially bottom end of the pressure side cavity, at least a first passage putting a radially top end of the suction side cavity into communication with a radially top end of the central cavity, at least a second passage putting a radially top end of the pressure side cavity into communication with the radially top end of the central cavity, and outlet orifices opening out both into the central cavity and into the pressure side face of the blade.

Abstract: A gas turbine blade for an airplane engine, the blade comprising at least a first cooling circuit comprising at least a concave side cavity extending radially beside the concave face of the blade, at least a second cooling circuit comprising at least one convex side cavity extending radially beside the convex face of the blade, and at least one third cooling circuit comprising at least one central cavity situated in the central portion of the blade between the concave side cavity and the convex side cavity, at least one leading edge cavity situated in the vicinity of the leading edge of the blade, communication orifices opening out into the central cavity and into the leading edge cavity, and outlet orifices opening out into the leading edge cavity and through the leading edge of the blade.

Abstract: A gas turbine blade for an airplane engine, the blade comprising at least a first cooling circuit comprising at least a concave side cavity extending radially beside the concave face of the blade, at least a second cooling circuit comprising at least one convex side cavity extending radially beside the convex face of the blade, and at least one third cooling circuit comprising at least one central cavity situated in the central portion of the blade between the concave side cavity and the convex side cavity, at least one leading edge cavity situated in the vicinity of the leading edge of the blade, communication orifices opening out into the central cavity and into the leading edge cavity, and outlet orifices opening out into the leading edge cavity and through the leading edge of the blade.

Abstract: A gas turbine blade for an airplane engine, comprising in its central portion at least a first central cooling circuit comprising at least first and second cavities extending radially on the concave side of the blade, at least one cavity extending on the convex side of the blade, an air admission opening at a radial end of the first concave side cavity for feeding the first 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 the concave face of the blade.