초록 ▼

A method enables a gas turbine engine blade to be manufactured to include an airfoil, a platform, a shank, and a dovetail. The platform extends between the airfoil and the shank, the shank extends between the dovetail and the platform, and the dovetail includes at least one tang for securing the bla

A method enables a gas turbine engine blade to be manufactured to include an airfoil, a platform, a shank, and a dovetail. The platform extends between the airfoil and the shank, the shank extends between the dovetail and the platform, and the dovetail includes at least one tang for securing the blade within the engine. The method comprises defining a cooling cavity in the blade that extends through the airfoil, the platform, the shank, and the dovetail, such that a portion of the cavity defined within the dovetail includes a root passage portion having a first width, and a transition portion extending 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. The blade is then coated with an environmental resistive coating.

대표청구항 ▼

1. 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 fo

1. 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. 2. A method in accordance with claim 1 wherein said defining a cooling cavity further comprises defining the cooling cavity within the dovetail such that the root passage first width is substantially constant within the root passage, and such that the transition portion is tapered between the root passage and the portion of the cavity defined within the shank, such that a width of the transition portion is variable within the transition portion.3. A method in accordance with claim 2 wherein said defining a cooling cavity further comprises defining the cooling cavity such that an interface between the dovetail transition portion and the shank portion forms an arcuate shape that defines a portion of the cooling cavity.4. A blade for a gas turbine engine, said blade comprising:a platform; a shank extending from said platform; a dovetail extending between an end of said blade and said shank for mounting said blade within the gas turbine engine, said dovetail comprising at least one tang; an airfoil comprising a first sidewall and a second sidewall extending in radial span between said platform and a blade tip; and a cooling cavity defined within said blade by said dovetail, said shank, said platform, and said airfoil, said cooling cavity comprising a dovetail portion defined within said dovetail, a shank portion defined within said shank and said platform, and an airfoil portion defined within said airfoil, said shank portion coupled in flow communication between said airfoil portion and said dovetail portion, said dovetail portion comprising a root passage and a transition passage, said root passage comprising a first width, said shank portion comprising a second width larger than said root passage first width, said transition passage coupled between said root passage and said shank portion, said dovetail further comprises an inner surface defining said cooling cavity dovetail portion, said dovetail inner surface coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches, such that said cooling cavity dovetail portion facilitates reducing dovetail low cycle fatigue cracking. 5. A blade in accordance with claim 4 wherein said cooling cavity root passage first width measured between a pressure side of said cooling cavity and a suction side of said cooling cavity, said root passage first width substantially constant within said root passage.6. A blade in accordance with claim 4 wherein said cooling cavity shank passage second width measured between a pressure side of said cooling cavity and a suction side of said cooling cavity, an interface of said transition passage and said shank portion is arcuate.7. A blade in accordance with claim 6 wherein said cooling cavity interface facilitates reducing operating stresses induced within said blade dovetail.8. A gas turbine engine comprising a plurality of blades, each said blade comprising an airfoil, a shank, and a platform extending therebetween, each said blade further comprising a cooling cavity, and a dovetail comprising at least one tang and configured to secure each said blade within said engine, said shank extending between said platform and said dovetail, said cooling cavity defined within said airfoil, said platform, said shank, and said dovetail, said cooling cavity comprising a dovetail portion, a shank portion, and an airfoil portion coupled in flow communication, said cooling cavity dovetail portion comprising a root passage comprising a first width, and a transition passage, said cooling cavity shank portion comprising a second width that is larger than said root passage first width, said cooling cavity transition passage tapering between said root passage and said shank portion, at least a portion of said cooling cavity is coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches such that said cooling cavity facilitates reducing dovetail low cycle fatigue cracking.9. A gas turbine engine in accordance with claim 8 wherein said cooling cavity root passage first width measured between a pressure side and a suction side of said cooling cavity, said cooling cavity shank portion second width is measured between said cooling cavity pressure and suction sides, said root passage first width is substantially constant within said root passage.10. A gas turbine engine in accordance with claim 9 wherein an interface between said cooling cavity transition passage and said cooling cavity shank portion forms a radius.11. A gas turbine engine in accordance with claim 9 wherein said cooling cavity coated with an oxidation resistant environmental coating having a thickness configured to reduce low cycle fatigue of each said blade.12. A gas turbine engine in accordance with claim 9 wherein at least a portion of said cooling cavity dovetail portion is coated with an oxidation resistant environmental coating having a thickness greater than 0.001 inches.13. A gas turbine engine in accordance with claim 9 wherein each said cooling cavity facilitates reducing operating stresses induced within each said blade dovetail.