Patent ID: 7497664

Claim:
A method for fabricating a rotor blade for a gas turbine engine wherein the rotor blade includes an airfoil including a first sidewall and a second sidewall connected at a leading edge and at a trailing edge, a root portion at a zero percent radial span and a tip portion at a one hundred percent radial span, the airfoil having a radial span dependent chord length C, a respective maximum thickness T max , and a maximum thickness to chord length ratio (T max /C ratio), said method comprises: determining a blade geometry that facilitates reducing a vibratory stress of the blade; and casting the rotor blade such that a root portion is formed having a first T max /C ratio, a tip portion is formed having a second T max /C ratio, and a mid portion, extending between the root portion and the tip portion, is formed having a third T max /C ratio, the third T max /C ratio being less than the first T max /C ratio and less than the second T max /C ratio, wherein the trailing edge is tapered and has a first thickness at about zero percent of span and a second thickness at about seventy percent of span.