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Methods and apparatus for fabricating a rotor blade for a gas turbine engine are provided. The rotor blade includes an airfoil having a first sidewall and a second sidewall, connected at a leading edge and at a trailing edge. The method includes forming the airfoil portion bounded by a root portion

Methods and apparatus for fabricating a rotor blade for a gas turbine engine are provided. The rotor blade includes an airfoil having a first sidewall and a second sidewall, connected at a leading edge and at a trailing edge. The method includes forming the airfoil portion bounded by 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, and a maximum thickness to chord length ratio (Tmax/C ratio), forming the root portion having a first Tmax/C ratio, forming the tip portion having a second Tmax/C ratio, and forming a mid portion extending between a first radial span and a second radial span having a third Tmax/C ratio, the third Tmax/C ratio being less than the first Tmax/C ratio and the second Tmax/C ratio.

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What is claimed is: 1. 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

What is claimed is: 1. 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 Tmax, and a maximum thickness to chord length ratio (Tmax/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 Tmax/C ratio, a tip portion is formed having a second Tmax/C ratio, and a mid portion, extending between the root portion and the tip portion, is formed having a third Tmax/C ratio, the third Tmax/C ratio being less than the first Tmax/C ratio and less than the second Tmax/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. 2. A method in accordance with claim 1 wherein casting the rotor blade further comprises: forming the root portion such that the first Tmax/C ratio is greater than about 0.08; forming the tip portion such that the second Tmax/C ratio is greater than about 0.06; and forming the mid portion centered at about sixty percent span such that the third T.sub.max/C ratio is less than about 0.05. 3. A method in accordance with claim 1 further comprising forming the trailing edge having a thickness that is tapered such that a thickness of said trailing edge decreases from the second thickness at about seventy percent of span to a third thickness at about one hundred percent span. 4. A method in accordance with claim 1 further comprising forming the leading edge having a first thickness at about zero percent of span, the leading edge thickness is tapered such that a thickness of said leading edge increases to a second thickness at about one hundred percent of span. 5. A method in accordance with claim 1 further comprising forming the leading edge having a first thickness at about zero percent of span, the leading edge thickness is tapered such that a thickness of said leading edge decreases continuously to a second thickness at about one hundred percent of span. 6. A method in accordance with claim 1 further comprising forming the tip portion with a greater Tmax/C ratio than the mid portion such that stripe mode stresses are facilitated being distributed over the tip portion and the mid portion. 7. A method in accordance with claim 1 further comprising forming the tip portion with a greater Tmax/C ratio than the mid portion such that stripe mode stresses are facilitated being reduced proximate the tip portion. 8. An airfoil for a gas turbine engine, said airfoil comprising a radial span dependent chord length C, a respective maximum thickness T, and a maximum thickness to chord length ratio (Tmax/C ratio), said airfoil further comprising: a first sidewall; a second sidewall coupled to said first sidewall at a leading edge and at a trailing edge, said trailing edge tapered such that a thickness of said trailing edge increases from about zero percent span to about seventy percent span; a root portion comprising a first Tmax/C ratio; a tip portion comprising a second Tmax/C ratio; and a mid portion extending between said root portion and said tip portion, said mid portion comprising a third Tmax/C ratio that is less than the first Tmax/C ratio and the second Tmax/C ratio. 9. An airfoil in accordance with claim 8 wherein said first Tmax/C ratio is greater than about 0.08, said second Tmax/C ratio is greater than about 0.06, and said third Tmax/C ratio is less than about 0.05. 10. An airfoil in accordance with claim 8 wherein said trailing edge is tapered such that a thickness of said trailing edge decreases from about seventy percent span to about one hundred percent span. 11. An airfoil in accordance with claim 8 wherein said leading edge is tapered such that a thickness of said leading edge decreases from about zero percent span to about one hundred percent span. 12. An airfoil in accordance with claim 11 further comprising forming the leading edge having a thickness that continuously decreases from about zero percent span to about one hundred percent span. 13. An airfoil in accordance with claim 8 further comprising forming the tip portion with a greater Tmax/C ratio than the mid portion such that stripe mode stresses are facilitated being distributed over the tip portion and the mid portion. 14. An airfoil in accordance with claim 8 further comprising forming the tip portion with a greater Tmax/C ratio than the mid portion such that stripe mode stresses are facilitated being reduced proximate the tip portion. 15. A gas turbine engine comprising a plurality of rotor blades, each said rotor blade comprising an airfoil comprising a radial span dependent chord length C, a respective maximum thickness T, and a maximum thickness to chord length ratio (Tmax/C ratio), said airfoil comprising: a first sidewall; a second sidewall coupled to said first sidewall at a leading edge and at a trailing edge, said trailing edge comprises a first thickness at about zero percent span, a second thickness at about one hundred percent of span, and a maximum thickness at about seventy percent of span; a root portion at a zero percent radial span having a first Tmax/C ratio; a tip portion at a one hundred percent radial span having a second Tmax/C ratio; and a mid portion extending between said root portion and said tip portion having a third Tmax/C ratio, the third Tmax/C ratio that is less than the first Tmax/C ratio and the second Tmax/C ratio. 16. A gas turbine engine in accordance with claim 15 wherein said first Tmax/C ratio is greater than about 0.08, said second Tmax/C ratio is greater than about 0.06, and said third Tmax/C ratio is less than about 0.05. 17. A gas turbine engine in accordance with claim 15 wherein said leading edge comprises a thickness that continuously decreases from a first leading edge thickness at about zero percent of span to a second leading edge thickness at about one hundred percent of span. 18. An airfoil for a gas turbine engine, said airfoil comprising: a first sidewall extending between a root portion and a tip portion; and a second sidewall extending between said root portion and said tip portion, said second sidewall coupled to said first sidewall at a leading edge and at a trailing edge; said airfoil comprising a maximum thickness, a leading edge thickness, a midchord thickness, and a trailing edge thickness wherein said trailing edge thickness is greater than said leading edge thickness, wherein each of said thicknesses is measured between said first and said second sidewalls. 19. An airfoil in accordance with claim 18 wherein said trailing edge thickness is at least 10% greater than said leading edge thickness. 20. An airfoil in accordance with claim 19 wherein said trailing edge thickness is at least 50% greater than said leading edge thickness. 21. An airfoil in accordance with claim 20 wherein said trailing edge thickness is approximately 100% greater than said leading edge thickness. 22. An airfoil in accordance with claim 18 wherein said maximum thickness is approximately equal to said midchord thickness. 23. An airfoil in accordance with claim 18 wherein said maximum thickness is less than 150% greater than said leading edge thickness. 24. An airfoil in accordance with claim 18 wherein said maximum thickness is less than 25% greater than said trailing edge thickness. 25. An airfoil in accordance with claim 18 wherein said maximum thickness is approximately 0.048 inches, said leading edge thickness is approximately 0.019 inches, said midchord thickness is approximately 0.047 inches, and said trailing edge thickness is approximately 0.04 inches.