Abstract:
A rotating airfoil component equipped with one or more angel wings that inhibit the ingress of a hot working fluid into interior regions of a turbomachine in which the component is installed. The component includes an airfoil and a feature for mounting the component to enable rotation of the component within the turbomachine. An angel wing projects from the component to have a first surface facing the airfoil, an oppositely-disposed second surface facing the mounting feature, and at least one lateral surface therebetween. A thermal-insulating coating system is present on the first surface to inhibit heat transfer from the working fluid to the angel wing but not on the second or lateral surfaces so as not to inhibit heat transfer from the second and lateral surfaces of the angel wing.

Description:
BACKGROUND OF THE INVENTION 
     The present invention generally relates to structures subject to high stresses and high temperatures, such as rotating components of gas turbines and other turbomachinery. More particularly, this invention relates to a method of inhibiting heat transfer to angel wings of turbine buckets (blades) so as to reduce the temperature of the angel wings and/or reduce the cooling requirements of the angel wings. 
     Buckets (blades), nozzles (vanes), and other components located in the hot gas path within turbine sections of gas turbines are typically formed of nickel-, cobalt- or iron-base superalloys with desirable mechanical and environmental properties for turbine operating temperatures and conditions. Because the efficiency of a gas turbine is dependent on its operating temperatures, there is a demand for components that are capable of withstanding increasingly higher temperatures. As the maximum local temperature of a component approaches the melting temperature of its alloy, forced air cooling becomes necessary. For this reason, airfoils of gas turbine buckets and nozzles often require complex cooling schemes in which air is forced through internal cooling passages within the airfoil and then discharged through cooling holes at the airfoil surface. 
       FIG. 1  schematically represents an axial cross-section of a turbine section  10  of a land-based gas turbine engine. The turbine section  10  comprises multiple turbine stages, represented as the first and second stages immediately downstream of the combustor (not shown) of the turbine engine. Each stage of the turbine section  10  comprises an annular array of circumferentially-spaced buckets  12  (only one bucket  12  of each stage is represented in  FIG. 1 ) and a nozzle assembly  14  made up of an annular array of circumferentially-spaced vanes  16  (only one vane  16  of each stage is represented in  FIG. 1 ). The nozzle assemblies  14  and their vanes  16  are statically mounted within the turbine section  10 , whereas the buckets  12  are mounted on a rotating component, commonly referred to as a wheel  18 , of the gas turbine to enable rotation of the buckets  12  within the gas turbine and relative to the nozzle assemblies  14 . The vanes  16  define airfoils that extend between inner and outer platforms (or bands)  20  of the nozzle assemblies  14 . As represented in  FIG. 1 , each bucket  12  comprises an airfoil  24  extending from a shank  26  in a radially outward direction  22  of the turbine section  10 . The buckets  12  can be conventionally anchored to their respective wheels  18 , for example, with dovetails (not shown) formed on their shanks  16  and received in complementary slots defined in the circumference of each wheel  18 . The buckets  12  and nozzle assemblies  14  are directly subjected to the hot gas path  32  within the turbine section  10 . In particular, the airfoils  24  of the buckets  12  and the vanes  16  of the nozzle assemblies  14  are impinged by the hot combustion gases in the hot gas path  32  through the gas turbine. 
     Impingement of the bucket airfoils  24  and nozzle vanes  16  by the combustion gases results in upstream airfoil wakes and downstream airfoil bow waves, which tend to produce pressure wakes within the hot gas path  32  that cause hot combustion gases to be driven into trench cavities  34  between rows of buckets  12  and nozzle assemblies  14  and, from there, into wheelspace cavities  36  between the wheels  18 . To inhibit the ingress of hot combustion gases into the interior regions of the gas turbine, the buckets  12  are commonly equipped with extensions, referred to as angel wings  28 , that extend from the shank  26  into the trench cavities  34  in a direction corresponding to the axial direction of the turbine section  10 . As represented in  FIG. 1 , the angel wings  28  cooperate with lands  30  formed on the adjacent nozzle assemblies  14  to create a tortuous path that inhibits the flow of hot gases through the trench cavities  34 . Consequently, the angel wings  28  are directly exposed to the hot combustion gases ingested into the trench cavities  34  from the gas path  32 . Current practice is to supply the trench cavities  34  with a cooling air flow  33  obtained by air bled from the compressor section (not shown) of the engine for the purpose of keeping the angel wings  28  at temperatures that are sufficiently low to enable the angel wings  28  to meet their creep and fatigue life requirements. However, this purge flow is costly to the overall performance of a gas turbine engine, and therefore any reduction in the cooling air flow  33  needed to protect the angel wings  28  would be advantageous to turbine efficiency. 
     BRIEF DESCRIPTION OF THE INVENTION 
     The present invention provides a rotating airfoil component of a turbomachine, and particular a component equipped with one or more angel wings that serve to inhibit the ingress of a working fluid (for example, hot combustion gases of a gas turbine or steam of a steam turbine) into interior regions of the turbomachine. 
     According to a first aspect of the invention, the component includes an airfoil adapted for impingement by the working fluid of the turbomachine, and means for mounting the component to a rotating component of the turbomachine to enable rotation of the component within the turbomachine. At least one angel wing projects from the component and is adapted to inhibit flow of the working fluid from the airfoil toward the mounting means of the component. The angel wing has a first surface facing the airfoil, an oppositely-disposed second surface facing the mounting means, and at least one lateral surface therebetween. A thermal-insulating coating system is present on the first surface to inhibit heat transfer from the working fluid to the angel wing but not on the second or lateral surfaces so as not to inhibit heat transfer from the second and lateral surfaces of the angel wing. 
     A technical effect of the invention is the ability to thermally protect an angel wing so that its creep and fatigue life requirements can be met, while allowing for higher operating temperatures within a turbomachine and/or a reduction in cooling air flow used to protect the angel wings. 
     Other aspects and advantages of this invention will be better appreciated from the following detailed description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically represents an axial cross-section of a turbine section of a land-based gas turbine engine. 
         FIG. 2  is a detailed view of an angel wing of the type represented in  FIG. 1  and in accordance with the prior art. 
         FIGS. 3-6  are detailed views of angel wings of the types represented in  FIGS. 1 and 2 , but modified to further have a thermal-insulating coating system in accordance with embodiments of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The invention will be described in reference to the turbine section  10  schematically represented in  FIG. 1 . The previous discussion of  FIG. 1  is therefore applicable to the following discussion, which will focus primarily on aspects of the invention that differ from what was previously described in reference to  FIG. 1 . However, it should be understood that the invention is not limited to the turbine section  10  and its particular configuration represented in  FIG. 1 . In particular, the invention is not limited to the particular buckets  12  represented in  FIG. 1 , but is more generally applicable to rotating airfoil components of turbomachines, including but not limited to gas turbines, land-based gas turbine engines, aircraft gas turbine engines, and steam turbines. Furthermore, the invention is not limited to the particular configurations and numbers of the angel wings  28  and lands  30  represented in  FIG. 1 . 
     The buckets  12  and components of the nozzle assembly  14  shown in  FIG. 1  can be conventionally formed of nickel-, cobalt-, or iron-based superalloys of types suitable for use in gas turbines. Notable but nonlimiting examples include nickel-based superalloys such as GTD-111® (General Electric Co.), GTD-444® (General Electric Co.), IN-738, René™ N4 (General Electric Co.), René™ N5 (General Electric Co.), René™ 108 (General Electric Co.) and René™ N500 (General Electric Co.). The buckets  12  and vanes  16  may be formed as equiaxed, directionally solidified (DS), or single crystal (SX) castings to withstand the high temperatures and stresses to which they are subjected within a gas turbine engine. Melting and casting processes suitable for producing the buckets  12  and vanes  16  are well known and therefore will not be discussed here in any detail. 
     As previously described in reference to  FIG. 1 , in combination with the lands  30 , the angel wings  28  serve to inhibit the inward flow of hot combustion gases (working fluid) from the airfoils  24  of the buckets  12 , through the trench cavities  34  toward the retention features (typically the dovetails) by which the bucket  12  is anchored to the wheel  18 , and into the wheelspace cavities  36 .  FIG. 2  is a detailed view of an angel wing  28  of the type shown in  FIG. 1 . The angel wing  28  is typically integrally cast with the remainder of its bucket  12 , to which the angel wing  28  is joined through root blends  38  to reduce stress concentrations. As indicated in  FIG. 2 , the angel wing  28  has radially outward and inward surface  40  and  42  (in relation to the radial direction  22  of the turbine section  10 ) and oppositely-disposed lateral surfaces  44  (of which only one is visible in  FIG. 2 ) between the outward and inward surfaces  40  and  42 . The angel wing  28  terminates at an upturned distal tip  46  that projects from the outward surface  40 . As evident from  FIG. 1 , the distal tip  46  projects toward the airfoil  24  in the radially outward direction  22  of the turbine section  10 , such that the tip  46  may be subjected to any rub encounters with the land  30  with which the angel wing  28  cooperates. The presence of an upturned distal tip  46  is advantageous in that the tip  46  can be machined to more closely control the radial height of the angel wing  28  relative to the bucket retention features (dovetails), enabling a narrower gap to be maintained with the corresponding land  30 . However, the present invention is also applicable to angel wings whose distal tips are not upturned, i.e., a flat-topped angel wing whose radially outward surface  40   a  is indicated in phantom in  FIG. 2 . 
     As evident from  FIG. 1 , the angel wing  28  is essentially cantilevered into one of the trench cavities  34  in a direction roughly perpendicular to the radial direction  22  of the turbine section  10 , such that its outward surface  40  generally faces the bucket airfoil  24  and its inward surface  42  generally faces the bucket retention features, as well as the wheelspace cavity  36  beyond. With the location and orientation represented in  FIG. 1 , the angel wing  28  will typically be surrounded by a mixture of the hot combustion gases that enter from the hot gas path  32  and air from the cooling air flow  33  bled from the compressor, with the result that the angel wing  28  tends to be at a rather uniform temperature, though typically with some degree of thermal gradient in the radial direction. As higher operating temperatures are sought to improve the efficiency of the gas turbine, the temperature of the angel wing  28  can rise to levels that unacceptably reduce its creep and fatigue life properties. 
     The present invention seeks to reduce the temperature of the angel wing  28  by reducing the heat flux into the angel wing  28  from the combustion gases ingested from the hot gas path  32 , and simultaneously take advantage of the high heat transfer coefficients and cooler air from the cooling air flow  33 . The invention does so by thermally insulating the radially outward surface  40  of the angel wing  28 , but not the inward or lateral surfaces  42  and  44  of the angel wing  28 , with a thermal-insulating coating system  48 , as represented in  FIG. 3 . In this manner, heat transfer from the hot combustion gases to the angel wing  28  can be reduced without reducing the ability of the angel wing  28  to transfer heat to the cooling air flow  33 . As represented in  FIG. 3 , in addition to being absent on the inward and lateral surfaces  42  and  44 , the coating system  48  is preferably not present on the distal tip  46  of the angel wing  28 , and therefore is not subjected to any rub encounters with the land  30 . However, as represented in  FIG. 4  at  48   a , it is also within the scope of the invention that the coating system  48  is deposited to also cover the radially outward surface of the distal tip  46 , whether configured to be upturned or flat-topped, such that the coating system  48  covers the entire radially outward surface of the angel wing  48 . The coating system  48  may also be deposited to fully wrap around only the distal tip  46  such that lateral and distal surfaces of the distal tip  46  are covered, as represented in  FIG. 5  at  48   b . Finally,  FIGS. 3-6  represent the coating system  48  as extending upward slightly over the radially-outward root blend  38  toward the airfoil  24 . 
     The coating system  48  that protects the radially outward surface  40  of the angel wing  28  can be of a type known in the art. Such systems, referred to as thermal barrier coating (TBC) systems, entail a low-conductivity thermal barrier coating (TBC) that is typically adhered to a substrate surface with a suitable bond coat. Typical but nonlimiting TBC materials for the coating system  48  are ceramic materials, a notable example of which is zirconia partially or fully stabilized with yttria (YSZ) or another oxide such as magnesia, ceria, scandia and/or calcia, and optionally other oxides to reduce thermal conductivity. A suitable thickness for the TBC is generally on the order of about 0.003 to about 0.050 inch (about 75 to about 1250 micrometers), with the upper limit intended to minimize the additional weight attributable to the TBC that could increase stresses in the angel wing  28 . Suitable techniques for depositing the TBC material include air plasma spraying (APS), suspension plasma spraying (SPS), electron beam physical vapor deposition (EB-PVD), plasma spray-physical vapor deposition (PS-PVD), etc. Masking can be used to prevent over-spray on surface regions that are not intended to be coated. 
     The coating system  48  preferably includes a metallic bond coat to promote the adhesion of the TBC material, whose ceramic composition results in a thermal expansion mismatch with the metallic composition of the angel wing  28 . Because the TBC material that provides a desired insulating effect may offer little resistance to oxidation, erosion, and corrosion, preferred bond coats are also capable of environmentally protecting the underlying outward surface  40  of the angel wing  28 . The ability of the bond coat to adhere the ceramic TBC and protect the underlying angel wing surface  40  can be promoted through the formation of an adherent oxide scale, such as a thin layer of aluminum oxide (alumina), on its surface, which chemically bonds the ceramic TBC to the bond coat. For this purpose, various bond coat materials have been proposed, notable examples of which have aluminum-rich compositions, including diffusion coatings that contain aluminum intermetallics (predominantly beta-phase nickel aluminide (β-NiAl) and platinum aluminides (PtAl)), and overlay coatings such as MCrAlX (where M is iron, cobalt and/or nickel, and X is yttrium, one or more rare earth metals, and/or one or more reactive metals), of which CoNiCrAlY and NiCrAlY are two notable examples. A suitable thickness for the bond coat is generally on the order of about 0.001 to about 0.015 inch (about 25 to about 380 micrometers). Suitable techniques for depositing the bond coat include APS, low pressure plasma spraying (LPPS, also referred to as vacuum plasma spaying, or VPS), high velocity air-fuel (HVAF) deposition, high velocity oxy-fuel (HVOF) deposition, ion plasma deposition (IPD, also called cathodic arc deposition), cold spraying, wire arc spraying, plating, etc. 
     By applying the coating system  48  to only the outward surface  40  of the angel wing  28 , and not to its inward or lateral surfaces  42  and  44 , the angel wing  28  is partially insulated from the hot combustion gases that are ingested from the hot gas path  32  and predominately encountered by the outward surface  40 . As such, the coating system  48  reduces the heat flux  50  into the angel wing  28  that would otherwise occur through its outward surface  40 , but does not insulate the inward and lateral surfaces  42  and  44  that predominantly encounter the cooling air flow  33  and therefore does not provide a barrier to heat transfer from these surfaces  42  and  44  to the cooling air flow  33 . Analytical investigations have indicated that locating the coating system  48  solely on the outward surface  40  of the angel wing  28  is capable of reducing the nominal temperature of an angel wing by roughly one hundred degrees Celsius or more. Consequently, it may be possible to increase the operating temperature of an engine corresponding to higher combustion gas temperatures) or reduce the cooling air flow  33  (corresponding to higher temperatures within the wheelspace cavities  36 ) while maintaining the angel wing  28  at a temperature compatible with meeting the required creep and fatigue life properties of the angel wing  28 . 
     While the invention has been described in terms of particular embodiments, it is apparent that other forms could be adopted by one skilled in the art. Therefore, the scope of the invention is to be limited only by the following claims.