| 摘要 |
The blade of gas turbine engine compressor of the invention includes a hammer type root to be inserted into a circumferential groove of the rotor of the compressor, a platform ( 12 ) integral with the root ( 11 ) and supporting an airfoil portion ( 13 ), the platform including two edges perpendicular to the axis of the rotor ( 20, 21 ) and two curved flanks ( 22, 23 ), the curve of the flanks being made out of at least one curve defined by an equation, the curvature centre of the point in the curve whereof the curvature radius is the smallest, being situated within a band (B) central to the platform and accounting for 60% of the width (D) of the platform ( 12 ) measured between its parallel rectilinear edges ( 20, 21 ), the equation defining the curve being, in the sense of mathematic functions, continuous and with continuous first derivative.
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