Gas turbine blade configuration
A gas turbine engine blade ( 22 ), including an airfoil substrate ( 10 ) having an exterior surface, wherein: a base ( 14 ) of the airfoil substrate is located at a 0% radial on an inner platform surface ( 20 ) and a tip ( 16 ) of the airfoil substrate is located at a 100% radial; wherein at the 0% radial a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 1; and wherein at a 50% radial location a cross-sectional profile of the exterior surface is characterized by nominal X and Y coordinates present in Table 6.
1. A gas turbine engine blade, comprising:
an airfoil substrate comprising an exterior surface, wherein a base of the airfoil substrate is located at a 0% radial location on an inner platform surface and a tip of the airfoil substrate is located at a 100% radial location,
wherein at the 0% radial a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 1, and
wherein at a 50% radial location a cross-sectional profile of the exterior surface is characterized by nominal X and Y coordinates present in Table 6.
2. The gas turbine engine blade of claim 1 , wherein at the 100% radial location a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 11.
3. The gas turbine engine blade of claim 2 , wherein at 10%, 20%, 30%, 40%, 60%, 70%, 80%, and 90% radial locations respective cross sectional profiles of the exterior surface are substantially characterized by nominal X and Y coordinates present in Tables 2, 3, 4, 5, 7, 8, 9, and 10 respectively.
4. The gas turbine engine blade of claim 1 , wherein the nominal X and Y coordinates represent dimensions in inches.
5. The gas turbine engine blade of claim 1 , further comprising a tip film cooling arrangement comprising an array of film cooling holes disposed on a pressure side of the airfoil substrate proximate the tip of the airfoil substrate.
6. The gas turbine engine blade of claim 5 , wherein the tip film cooling holes comprise a 10-10-10 shape angle orientation.
7. The gas turbine engine blade of claim 1 , further comprising a bond coat disposed on the airfoil substrate, and a thermal barrier coating disposed on the bond coat.
8. A gas turbine engine comprising a turbine, wherein a first stage of the turbine comprises the gas turbine engine blade of claim 1 .
9. A gas turbine engine blade, comprising:
an airfoil substrate comprising an exterior surface, wherein a base of the airfoil substrate is located at a 0% radial location on an inner platform and a tip of the airfoil substrate is located at a 100% radial location,
wherein at the 0% radial location a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 1, and wherein a lowest nominal X value in Table 1 defines a 0% radial leading edge point and a 0% radial leading edge point nominal Y value;
wherein at a 50% radial location a cross-sectional profile of the exterior surface comprises a 50% radial leading edge point characterized by a lowest nominal X value in Table 6.
10. The gas turbine engine blade of claim 9 , wherein at the 100% radial location radial a cross-sectional profile of the exterior surface comprises a 100% radial leading edge point characterized by a lowest nominal X value in Table 11.
11. The gas turbine engine blade of claim 9 , wherein the nominal X and Y coordinates represent dimensions in inches.
12. The gas turbine engine blade of claim 9 , further comprising a tip film cooling arrangement comprising an array of film cooling holes disposed on a pressure side of the airfoil substrate proximate the tip of the airfoil substrate.
13. The gas turbine engine blade of claim 9 , further comprising a bond coat disposed on the airfoil substrate, and a thermal barrier coating disposed on the bond coat.
14. A gas turbine engine blade, comprising:
an airfoil comprising an exterior surface, wherein a base of the airfoil is located at a 0% radial location on an inner platform surface and a tip of the airfoil is located at a 100% radial location,
wherein at the 0% radial location a cross-sectional profile of the exterior surface lies within a 0% radial envelope based on nominal X and Y coordinates present in Table 1,
wherein at a 50% radial location a cross-sectional profile of the exterior surface lies within a 50% radial envelope based on nominal X and Y coordinates present in Table 6,
wherein respective envelopes are defined by a respective nominal profile connecting respective nominal X and Y coordinates, minus an maximum inward variation of 0.015 inches inward from the respective nominal profile in a direction normal to the surface at that location, and plus a maximum outward variation of 0.060 inches outward from the respective nominal profile in a direction normal to the surface at that location.
15. The gas turbine engine blade of claim 14 , wherein at the 100% radial location a cross-sectional profile of the exterior surface lies within a 100% radial envelope based on nominal X and Y coordinates present in Table 11, and
wherein the 100% radial envelope is defined by a nominal 100% radial profile connecting respective nominal X and Y coordinates, minus an maximum inward variation of 0.015 inches inward from the nominal 100% radial profile in a direction normal to the surface at that location, and plus a maximum outward variation of 0.060 inches outward from the nominal 100% radial profile in a direction normal to the surface at that location.
16. The gas turbine engine blade of claim 15 , wherein at 10%, 20%, 30%, 40%, 60%, 70%, 80%, and 90% radial locations respective cross sectional profiles of the exterior surfaces lie within radial envelopes based on nominal X and Y coordinates present in Tables 2, 3, 4, 5, 7, 8, 9, and 10 respectively, and
wherein respective envelopes are defined by a respective nominal profile connecting respective nominal X and Y coordinates, minus an maximum inward variation of 0.015 inches inward from the respective nominal profile in a direction normal to the surface at that location, and plus a maximum outward variation of 0.060 inches outward from the respective nominal profile in a direction normal to the surface at that location.
17. The gas turbine engine blade of claim 14 , wherein the airfoil consists of a casting.
18. The gas turbine engine blade of claim 14 , further comprising a bond coat disposed on an airfoil substrate.
19. The gas turbine engine blade of claim 18 , further comprising a TBC disposed on the bond coat.
20. The gas turbine engine blade of claim 14 , further comprising a tip film cooling arrangement comprising an array of film cooling holes disposed on a pressure side of the airfoil proximate the tip of the airfoil.