| 摘要 |
PROBLEM TO BE SOLVED: To improve a gas turbine blade for an aircraft engine and provide a cooling circuit of the blade. SOLUTION: The blade 1 comprises: at least a first cooling circuit A provided with at least one recess side cavity 2 extending, near a recess surface 1a of the blade, in a radial direction; at least a second cooling circuit B separated from the first cooling circuit A and provided with at least one projection side cavity 4 extending, near a projection surface of the blade, in the radial direction; and at least a third cooling circuit C separated from the first and second cooling circuits. The third cooling circuit C comprises: at least one center cavity 6 arranged to a center part between the recess side cavity 2 and the projection side cavity 4, of the blade; at least one leading edge cavity 8 arranged near a front end 1c of the blade; a communication orifice opening to the center cavity and the leading edge cavity; and an outlet orifice opening to an inside of the leading edge cavity via the front end 1c of the blade.
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