Patent ID: 6932570

Claim:
A method for manufacturing a blade for a gas turbine engine, wherein the blade includes an airfoil, a platform, a shank, and a dovetail, the platform extending between the airfoil and the shank, the shank extending between the dovetail and the platform, the dovetail including at least one tang for securing the blade within the engine, said method comprising: defining a cooling cavity in the blade that extends through the airfoil the platform, the shank, and the dovetail, wherein the portion of the cavity defined within the dovetail includes a root passage portion having a first width, and a transition portion that extends between the root passage and the portion of the cavity defined within the shank, and wherein the portion of the cavity defined within the shank has a second width that is larger than the root passage first width; and coating at least a portion of an inner surface of the blade that defines the cooling cavity with a layer of an oxidation resistant environmental coating, such that at least a portion of the inner surface of the cooling cavity within the dovetail is coated with a coating having a thickness greater than 0.001 inches to facilitate reducing life cycle fatigue cracking within the dovetail.