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.
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0. 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.
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.
0. 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.
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 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;
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.
0. 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.
0. 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.
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
4. The airfoil turbine nozzle of
5. The airfoil turbine nozzle of
7. The airfoil turbine nozzle of
8. The airfoil turbine nozzle of
9. The airfoil turbine nozzle of
10. The airfoil turbine nozzle of
11. The airfoil turbine nozzle of
12. The airfoil turbine nozzle of
13. The airfoil turbine nozzle of
14. The airfoil turbine nozzle of
17. The airfoil nozzle assembly of
18. The airfoil nozzle assembly of
19. The airfoil nozzle assembly of
20. The airfoil nozzle assembly of
21. The airfoil nozzle assembly of
22. The airfoil nozzle assembly of
23. The airfoil nozzle assembly of
24. The airfoil nozzle assembly of
25. The airfoil nozzle assembly of
0. 27. The method of
0. 28. The method of
0. 29. The method of
0. 32. The turbine nozzle of
0. 33. The turbine nozzle of
0. 34. The turbine nozzle of
0. 35. The turbine nozzle of
0. 36. The turbine nozzle of
0. 37. The turbine nozzle of
0. 38. The turbine nozzle of
0. 39. The turbine nozzle of
0. 40. The turbine nozzle of
0. 41. The turbine nozzle of
0. 42. The turbine nozzle of
0. 43. The turbine nozzle of
0. 45. A turbine nozzle in accordance with
0. 46. A turbine nozzle in accordance with
0. 47. A turbine nozzle in accordance with
0. 48. A turbine nozzle in accordance with
0. 49. A turbine nozzle in accordance with
0. 50. A turbine nozzle in accordance with
0. 51. A turbine nozzle in accordance with
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1. Field of the Invention
This invention relates to thermal barrier coating systems for airfoils used in the turbine portions of gas turbine engines, and specifically to an improved coating arrangement for segments consisting of a plurality of airfoils.
2. Discussion of the Prior Art
blades used with blade assemblies, which rotate, or vanes vane assemblies, which are fixed in position in relation to a gas turbine engine. Blades are attached to rotors through a base, thereby rotating with the rotor. The base is not considered part of the flow path surface.
Referring to
Referring now to
Because of the differences in mass between outer band 14 and body 16, a certain amount of stresses are introduced to blades airfoils at the juncture simply of the outer band and the airfoil body as a matter of thermal stresses resulting from heat transfer from a large mass to a much smaller mass. Additive to these stresses are the mechanical stresses due to the effects of the transition radius 30. Finally, additional stresses resulted from the effects of the termination of the thermal barrier coating (TBC) in the region or the transition radius.
The present invention reduces these stresses from the effects of terminating the thermal coating in the region of the transition radius by extending the coating below radius 30 for a distance sufficient to reduce the thermal stresses in the region of radius 30 resulting from the abrupt termination of the TBC in the region of radius 30, thereby reducing the overall accumulation of stresses in this region. This distance is indicated by d in FIG. 2. It is preferred that this distance d extend from the fillet radius in the direction of the base such that at least about 20% of the span between the radius and the base is coated. It is preferred that d be at least about 0.5″ and preferably in the range of 0.5″ to 0.6″ for an airfoil such as the one found on the CF6 engine, but larger or smaller values are acceptable as long as the stresses resulting from the transition are lowered. Since nozzle segment 10 is comprised of two airfoils 12, the application of the bond coat and TBC is limited by the ability of the spray nozzles used for their applications to effectively apply coating between the airfoils 12. Thus, the TBC is applied to at least a portion of body 16 from line 36 bisecting the airfoil to the trailing edge for a distance d so that the thermal barrier coating applied to the body blends smoothly with the preexisting thermal barrier coating covering the leading edge region so that the TBC is extended from the line bisecting the airfoil toward the trailing edge. For airfoils such as those found in a CF6 application, this coating will extend from the leading edge region toward the trailing edge about at least another 0.9″. Preferably, the TBC is applied to both sides of each airfoil, but the ability of the nozzle to apply material may limit the application of material to the portions of the airfoils which face each other. The bond coat and the ceramic top coat are both applied using air plasma spray and line-of-sight application techniques. In addition, the portion of the trailing edge on the concave or pressure side includes cooling holes. Because it is unacceptable to have these holes adversely impacted by coating, it is necessary to mask these holes to prevent blockage by the TBC. This is an acceptable compromise since the cooling film resulting from the air flow assists in keeping this region cool.
While the airfoils may be masked so that a TBC may be applied to full thickness, it is preferred that the TBC be tapered from full thickness, 0.012-015″ 0.012-0.015″ at radius 30 to about 0.002-0.005 inches at a preferred distance d of 0.5-0.6″. Thus, assuming a linear taper, the preferred slope of the TBC thickness is in the range of about 0.020-0.150 in./in. and most preferably 0.021-0.140 in./in. This further reduces the possibility of thermally-induced stresses in the vicinity of the termination of the TBC.
The thermal barrier coating system of the present invention is also applied in a manner somewhat different than the prior art thermal barrier coating systems. First, the bond coat, a MCrAlY(X) where 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 is applied by an air plasma spray process. The segment is then aluminided using a typical vapor phase aluminiding procedure. Finally, a the ceramic topcoat, 7% Yttria-Stabilized Zirconia (7YSZ) is applied over the bond coat in the region set forth above by air plasma spraying, which is an improvement over the low pressure plasma spray methods previously used to apply ceramic topcoats.
Referring now to
The TBC wrap of the present invention is a new concept for high pressure turbine applications. The use of this wrap as set forth above, by extending the TBC below fillet region 30 and onto the flow path surfaces of the airfoil body will result in decreased metal temperatures, a more balanced thermal design and reduced peak stresses. Whereas the prior art airfoil design experienced a temperature gradient of about 80-100° F. over a distance of about ¼″, the present invention lowers the thermal gradient to about 50-80° F. over a distance of about ½″. This decrease is expected to result in at least a 50% improvement in LCF life, which is a conservative estimate. Along with lower stresses and lower temperatures in this region, there will be reduced incidents of cracking and burning.
Although the present invention has been described in connection with specific examples and embodiments, those skilled in the art will recognize that the present invention is capable of other variations and modifications within its scope. These examples and embodiments are intended as typical of. of, rather than in any way limiting on, the scope of the present invention as presented in the appended claims.
Nichols, Glenn H., Tressler, Judd Dodge
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