Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the rectangular coordinates X,Y,Z? given in Table 1, in which the X coordinate is measured in the axial direction of the turbine, and the Z? coordinate is the quotient D/H, where D is the distance of the point in question from a reference plane P0 situated at the base of the nominal profile and H is the height of the profile measured from said reference plane to the tip of the blade, the measurements D and H being taken radially relative to the axis of the turbine.

Abstract: Optimized aerodynamic airfoil for a turbine blade When cold and in the uncoated state, the aerodynamic airfoil is substantially identical to a nominal airfoil determined by the Cartesian coordinates X, Y, Z? given in Table 1, in which the coordinate Z? is the ratio D/H where D is the distance from the point in question to a reference plane P0, located at the base of the nominal airfoil, and H is the height of this airfoil, measured from said reference plane up to the tip of the blade. The D and H measurements being taken radially with respect to the axis of the turbine whereas the X coordinate is measured in the axial direction of the turbine.

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the rectangular coordinates X,Y,Z? given in Table 1, in which the X coordinate is measured in the axial direction of the turbine, and the Z? coordinate is the quotient D/H, where D is the distance of the point in question from a reference plane P0 situated at the base of the nominal profile and H is the height of the profile measured from said reference plane to the tip of the blade, the measurements D and H being taken radially relative to the axis of the turbine.

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 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.

Abstract: A vacuum switch with an elongated cartridge comprising a cylindrical housing sealed by two end plates. Inside the cylindrical housing there are two arcing contacts, one stationary and one movable. The stationary contact is securedly affixed to one of the end plates, whereas the other movable contact is mounted to axially slide inside the cartridge. At least two portions of turn are mounted in parallel. Each portion has a first end electrically connected to one of the arcing contacts and a second end electrically connected to the current input strip of the contact. The first and second ends of each portion of turn are electrically connected by a branch-off designed to disperse a part of the main current trough the potions of turn during breaking.