Thermal barrier coating wrap for turbine airfoil

Abstract

An airfoil having extended life due to reduction in stresses. The stresses are reduced by extending the thermal barrier coating below the radius between the outer band perimeter and the inner flow path surfaces and tapering the coating thickness. This additional tapered thermal barrier coating reduces the temperature gradient across a region already having high mechanical stresses resulting from geometric considerations thereby lowering thermally-induced stresses so that low cycle fatigue life is improved.

1. An improved airfoil, the turbine nozzle comprising an airfoil having flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edge edges opposite the pressure side of the airfoil, an outer band perimeter that includes a tip portion and an outer band portion of the airfoil , and a fillet radius extending from the outer band portion perimeter, the fillet radius forming a smooth contour between the outer band portion perimeter and the leading edge, the trailing edge, the concave side of the blade airfoil and the convex side of the blade airfoil, the improvement comprising:
 a thermal barrier coating system that extends over a preselected region of the flow path surface that includes the outer band perimeter, the fillet radius and at least a portion of the flow path surface on at least one side of the airfoil between the trailing edge and about a midpoint between the trailing edge and the leading edge of the airfoil for a distance below the fillet radius sufficient to reduce cumulative stresses resulting from a combination of mechanically-induced stresses from the fillet radius and service-induced thermal stresses, the thermal barrier coating system including a bond coat applied over the preselected region of the airfoil substrate, an environmental aluminiding coating applied over at least the bond coat and a ceramic coating having low thermal conductivity applied over the bond coat and the aluminide coating.  
 
     
     
       2. An airfoil A turbine nozzle for service in the turbine portion of a gas turbine engine, said turbine nozzle comprising at least one airfoil comprised of:
 flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edge edges opposite the pressure side of the airfoil and airfoil, an outer band perimeter that includes a tip portion and an outer band portion of the airfoil, and a fillet radius extending from the outer band portion, the fillet radius forming a smooth contour between the outer band portion and the leading edge, the trailing edge, the concave side of the airfoil and the convex side of the airfoil; and  
 a base integrally attached to the airfoil opposite the outer band perimeter;  
 the airfoil having a thermal barrier coating system that coats the fillet radius, the outer band perimeter portion and a leading edge region of the flow path surfaces, the leading edge region including the concave side and the convex side of the airfoil from the leading edge to about a line bisecting the airfoil between the leading edge and the trailing edge and spanning the airfoil from the fillet radius to the base, and a preselected region including at least a portion of the flow path surface on at least one side of the airfoil between the trailing edge and about a line bisecting the airfoil between the trailing edge and the leading edge for a preselected distance below the fillet radius toward the base sufficient to reduce cumulative stresses in a region of the fillet radius resulting from a combination of mechanically-induced stresses from the fillet radius and service-induced thermal stresses in the region, the thermal barrier coating system being continuous with the thermal barrier coating system in the leading edge region for the preselected distance, the thermal barrier coating system including a bond coat applied over the preselected region of the airfoil substrate, an environmental aluminiding coating applied over at least the bond coat and a ceramic coating having low thermal conductivity applied over the bond coat and the aluminide coating.  
 
     
     
       3. The airfoil turbine nozzle of  claim 2  wherein the thermal barrier coating system that extends over the preselected region of the airfoil has a maximum coating thickness at the outer band perimeter, the thickness of the coating decreasing at a preselected rate along the preselected distance from the fillet radius toward the base to reduce stresses from service-related thermal gradients. 
     
     
       4. The airfoil turbine nozzle of  claim 2  wherein the preselected distance over which the thermal barrier coating extends is at least about 20% of a span of the airfoil between the fillet radius and the base. 
     
     
       5. The airfoil turbine nozzle of  claim 2  wherein the preselected distance over which the thermal barrier coating extends is at least about 0.5 inches. 
     
     
       6. The airfoil turbine nozzle of  claim 2  wherein the preselected distance is about 0.5-0.6 inches. 
     
     
       7. The airfoil turbine nozzle of  claim 2  wherein the preselected region on a portion of the flow path surface is located on the convex side of the airfoil. 
     
     
       8. The airfoil turbine nozzle of  claim 2  wherein the preselected region on a portion of the flow path surface is located on the concave side of the airfoil. 
     
     
       9. The airfoil turbine nozzle of  claim 2  wherein the thermal barrier coating system coats at least the preselected region for a distance extending at least about 0.9 inches from the coated leading edge region toward the trailing edge. 
     
     
       10. The airfoil turbine nozzle of  claim 3  wherein the preselected rate is about 0.020 to about 0.15 inches per inch. 
     
     
       11. The airfoil turbine nozzle of  claim 2  wherein the thermal barrier coating system that extends over the preselected region of the airfoil has a substantially constant coating thickness from fillet radius toward the base to reduce stresses from service-related thermal gradients, the thermal barrier coating terminating at the preselected distance. 
     
     
       12. The airfoil turbine nozzle of  claim 2  wherein the bond coat is applied by air plasma spraying along a line-of-sight extending from the trailing edge toward the leading edge. 
     
     
       13. The airfoil turbine nozzle of  claim 2  wherein the ceramic coating is applied by air plasma spraying along a line-of-sight extending from the trailing edge toward the leading edge. 
     
     
       14. The airfoil turbine nozzle of  claim 2  wherein the bond coat is an MCrAlY MCrAlY( X ) in which M is at least one element selected from the group consisting of Ni, Co, and Fe and X is at least one element selected from the group consisting of Ti, Ta, Ru, Pt, Si, B, C, Hf and Zr. 
     
     
       15. The airfoil turbine nozzle of  claim 2  wherein the ceramic coating is YSZ. 
     
     
       16. An airfoil A nozzle assembly for service in the turbine portion of a gas turbine engine, said nozzle assembly comprised of:
 at least one airfoil comprising flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edge opposite the pressure side of the airfoil andairfoil, an outer band perimeter that includes a tip portion and an outer band portion of the airfoil, and a fillet radius extending from the outer band portion, the fillet radius forming a smooth contour between the outer band portion and the leading edge, the trailing edge, the concave side of the airfoil and the convex side of the airfoil;  
 a base integrally attached to the airfoil opposite the outer band perimeter;  
 the airfoil having a thermal barrier coating system that coats the outer band perimeter portion, the fillet radius radius, and a region extending from the leading edge to at least about ⅔ of the distance to the trailing edge, and from the fillet radius for a preselected distance below the fillet radius toward the base sufficient to reduce cumulative stresses in a region of the fillet radius resulting from a combination of mechanically-induced stresses from the fillet radius and service-induced thermal stresses in the region, the thermal barrier coating system including a bond coat, an environmental aluminiding coating applied over at least the bond coat and a ceramic coating having low thermal conductivity applied over the bond coat and the aluminide coating.  
 
     
     
       17. The airfoil nozzle assembly of  claim 16  wherein the thermal barrier coating system that extends over the preselected region of the airfoil has a maximum coating thickness at the outer band perimeter portion, the thickness of the coating decreasing at a preselected rate along the preselected distance from the fillet radius toward the base to reduce stresses from service-related thermal gradients. 
     
     
       18. The airfoil nozzle assembly of  claim 16  wherein the preselected distance over which the thermal barrier coating extends is at least about 20% of a span of the airfoil between the fillet radius and the base. 
     
     
       19. The airfoil nozzle assembly of  claim 16  wherein the preselected distance over which the thermal barrier coating extends is at least about 0.5 inches. 
     
     
       20. The airfoil nozzle assembly of  claim 16  wherein the preselected distance is about 0.5-0.6 inches. 
     
     
       21. The airfoil nozzle assembly of  claim 16  wherein the region extending from the leading edge is located on the convex side of the airfoil. 
     
     
       22. The airfoil nozzle assembly of  claim 16  wherein the region extending from the leading edge is located on the concave side of the airfoil. 
     
     
       23. The airfoil nozzle assembly of  claim 16  wherein the thermal barrier coating system coats from the leading edge at least about 1.8 inches toward the trailing edge. 
     
     
       24. The airfoil nozzle assembly of  claim 17  wherein the preselected rate is about 0.020 to about 0.15 inches per inch. 
     
     
       25. The airfoil nozzle assembly of  claim 16  wherein the thermal barrier coating system has a substantially constant coating thickness from fillet radius toward the base to reduce stresses from service-related thermal gradients, the thermal barrier coating terminating at the preselected distance. 
     
     
       26. A method to reduce cumulative stresses resulting from a combination of mechanically- induced stresses from the fillet radius and service - induced thermal stresses in an airfoil, comprising the steps of:      providing a nozzle assembly having airfoil flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edges opposite the pressure side of the airfoil, and also having an outer band perimeter that includes a tip portion and an outer band portion of the airfoil, and a fillet radius extending from the outer band portion, the fillet radius forming a smooth contour between the outer band portion and the leading edge, the trailing edge, the concave side of the airfoil and the convex side of the airfoil;        applying a bond coat along the leading edge from the outer band portion of the airfoil to a base of the airfoil;        applying a bond coat in a leading edge region defined by the fillet radius, the outer band portion and a leading edge region of the flow path surfaces, the leading edge region including the concave side and the convex side of the airfoil from the leading edge to about a line bisecting the airfoil between the leading edge and the trailing edge and spanning the airfoil from the fillet radius to the base;        applying a bond coat to a preselected region including at least a portion of the flow path surface on at least one side of the airfoil between the trailing edge and about a line bisecting the airfoil between the trailing edge and the leading edge for a preselected distance below the fillet radius toward the base sufficient to reduce cumulative stresses in a region of the fillet radius, the bond coat applied to the preselected region being continuous with the bond coat in the leading edge region for the preselected distance;        applying a ceramic topcoat over the bond coat in the leading edge region and the preselected region.     
     
     
       27. The method of  claim 26  wherein the step of applying the bond coat is by an air plasma spray method. 
     
     
       28. The method of  claim 26  wherein the step of applying the ceramic top coat is by an air plasma spray method. 
     
     
       29. The method of  claim 26  further including an additional step of masking cooling holes in the airfoil prior to application of the ceramic top coat. 
     
     
       30. A method to reduce cumulative stresses resulting from a combination of mechanically- induced stresses from the fillet radius and service - induced thermal stresses in an airfoil, comprising the steps of:      providing a nozzle assembly having airfoil flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edges opposite the pressure side of the airfoil, an outer band perimeter that includes a tip portion and an outer band portion of the airfoil, and a fillet radius extending from the outer band portion, the fillet radius forming a smooth contour between the outer band portion and the leading edge, the trailing edge, the concave side of the airfoil and the convex side of the airfoil, the airfoil also having a thermal barrier coating applied to the leading edge;        applying a bond coat to a preselected region including at least a portion of the flow path surface on at least one side of the airfoil between the trailing edge and about a line bisecting the airfoil between the trailing edge and the leading edge for a preselected distance below the fillet radius toward the base sufficient to reduce cumulative stresses in a region of the fillet radius, the bond coat applied to the preselected region being continuous with the thermal barrier coating applied to the leading edge region for the preselected distance;        applying a ceramic topcoat over the bond coat in the preselected region.     
     
     
       31. A turbine nozzle for a gas turbine engine, said turbine nozzle comprising:
   at least one airfoil comprising flow path surfaces that include a leading edge, a trailing edge, a concave side extending between the leading and trailing edges on a pressure side of the airfoil, and a convex side extending between the leading and trailing edges opposite the pressure side of the airfoil; and        an outer band perimeter that includes a tip portion and an outer band portion of the airfoil;        a fillet radius extending from the outer band portion, the fillet radius forming a smooth contour between the outer band portion and the leading edge, the trailing edge, the concave side of the airfoil and the convex side of the airfoil; and        a base integrally attached to the airfoil opposite the outer band perimeter;        the airfoil having a thermal barrier coating system that coats the fillet radius, a portion of the outer band perimeter and a leading edge region of the flow path surfaces, the leading edge region including the concave side and the convex side of the airfoil from the leading edge to about a line bisecting the airfoil between the leading edge and the trailing edge and spanning the airfoil from the fillet radius to the base, and a preselected region including at least a portion of the flow path surface on at least one side of the airfoil between the trailing edge and about a line bisecting the airfoil between the trailing edge and the leading edge for a preselected distance below the fillet radius toward the base sufficient to reduce cumulative stresses in a region of the fillet radius resulting from a combination of mechanically - induced stresses from the fillet radius and service - induced thermal stresses in the region, the thermal barrier coating system being continuous with the thermal barrier coating system in the leading edge region for the preselected distance, the thermal barrier coating system including a bond coat applied over the preselected region of the airfoil substrate, and a ceramic coating having low thermal conductivity applied over the bond coat.     
     
     
       32. The turbine nozzle of  claim 31  wherein the thermal barrier coating system that extends over the preselected region of the airfoil has a maximum coating thickness at the outer band perimeter, the thickness of the coating decreasing at a preselected rate along the preselected distance from the fillet radius toward the base to reduce stresses from service- related thermal gradients.   
     
     
       33. The turbine nozzle of  claim 31  wherein the preselected distance over which the thermal barrier coating extends is at least about  20 %  of a span of the airfoil between the fillet radius and the base.   
     
     
       34. The turbine nozzle of  claim 31  wherein the preselected distance over which the thermal barrier coating extends is at least about  0 . 5  inches. 
     
     
       35. The turbine nozzle of  claim 31  wherein the preselected distance is about  0 . 5 - 0 . 6  inches. 
     
     
       36. The turbine nozzle of  claim 31  wherein the preselected region on a portion of the flow path surface is located on the convex side of the airfoil. 
     
     
       37. The turbine nozzle of  claim 31  wherein the preselected region on a portion of the flow path surface is located on the concave side of airfoil. 
     
     
       38. The turbine nozzle of  claim 31  wherein the thermal barrier coating system coats at least the preselected region for a distance extending at least about  0 . 9  inches from the coated leading edge region toward the trailing edge. 
     
     
       39. The turbine nozzle of  claim 31  wherein the thermal barrier coating system that extends over the preselected region of the airfoil has a substantially constant coating thickness from fillet radius toward the base to reduce stresses from service- related thermal gradients, the thermal barrier coating terminating at the preselected distance.   
     
     
       40. The turbine nozzle of  claim 31  wherein the bond coat is applied by air plasma spraying along a line- of - sight extending from the trailing edge toward the leading edge.   
     
     
       41. The turbine nozzle of  claim 31  wherein the ceramic coating is applied by air plasma spraying along a line- of - sight extending from the trailing edge toward the leading edge.   
     
     
       42. The turbine nozzle of  claim 31  wherein the bond coat is an MCrAlY( X )  in which M is at least one element selected from the group consisting of Ni, Co, and Fe and X is at least one element selected from the group consisting of Ti, Ta, Ru, Pt, Si, B, C, Hf and Zr.   
     
     
       43. The turbine nozzle of  claim 31  wherein the ceramic coating is YSZ. 
     
     
       44. A turbine nozzle for a gas turbine engine, said nozzle comprising:
   a radially outer band;        at least one airfoil extending from said outer bands, said airfoil comprising a first sidewall and a second sidewall connected together at a leading edge and a trailing edge,        a fillet radius extending between said at least one airfoil and said outer band, said fillet radius extending between said airfoil leading and trailing edges and forming a substantially smooth contour between said at least one airfoil and said outer band; and        a thermal barrier coating extending over at least a portion of the radially outer band, the fillet radius, and a pre - selected region of said airfoil extending across at least one of said airfoil first and second sidewalls from said airfoil leading edge aft to at least a midpoint between said airfoil leading and trailing edges and for a distance below said fillet radius sufficient to facilitate reducing stresses induced to said airfoil, said thermal barrier coating has a variable thickness measured across said airfoil.     
     
     
       45. A turbine nozzle in accordance with  claim 44  wherein said thermal barrier coating comprises a bond coating applied over pre- selected region, an environmental aluminiding coating applied over at least said bond coating, and a ceramic coating applied over at least said aluminiding coating such that said aluminiding coating is between said ceramic and bond coatings.   
     
     
       46. A turbine nozzle in accordance with  claim 45  wherein said bond coating is an MCrAlY( X )  in which M is at least one element selected from the group consisting of Ni, Co, and Fe and X is at least one element selected from the group consisting of Ti, Ta, Ru, Pt, Si, B, C, Hf and Zr.   
     
     
       47. A turbine nozzle in accordance with  claim 45  wherein said ceramic coating is YSZ. 
     
     
       48. A turbine nozzle in accordance with  claim 44  wherein said thermal barrier coating is applied with a first thickness across said radially outer band and a second thickness across said airfoil, wherein said first thickness is thicker than said second thickness. 
     
     
       49. A turbine nozzle in accordance with  claim 48  wherein said thermal barrier coating thickness is tapered from said radially outer band through said airfoil, such that a thickness of thermal barrier coating applied to said radial outer band is thicker than a thickness of thermal barrier coating applied to said airfoil. 
     
     
       50. A turbine nozzle in accordance with  claim 44  wherein said airfoil first sidewall is convex and said airfoil second sidewall is concave, said thermal barrier coating extends at least partially across said first sidewall. 
     
     
       51. A turbine nozzle in accordance with  claim 44  wherein said airfoil first sidewall is convex and said airfoil second sidewall is concave, said thermal barrier coating extends at least partially across said second sidewall.