The present disclosure is directed to a method for designing the surface geometry associated with geometrically segmented abradable ceramic (GSAC) thermal barrier coating (TBC) on parts with complex shape.
Components that are exposed to high temperatures, such as a component within a gas turbine engine, typically include protective coatings. For example, components such as turbine blades, turbine vanes, blade outer air seals, and compressor components typically include one or more coating layers that function to protect the component from erosion, oxidation, corrosion or the like to thereby enhance component durability and maintain efficient operation of the engine.
As an example, some conventional turbine blade outer air seals include an abradable ceramic coating that contacts tips of the turbine blades such that the blades abrade the coating upon operation of the engine. The abrasion between the outer air seal and the blade tips provide a minimum clearance between these components such that gas flow around the tips of the blades is reduced to thereby maintain engine efficiency. Over time, internal stresses can develop in the protective coating to make the coating vulnerable to erosion and spalling. The outer air seal may then need to be replaced or refurbished after a period of use.
Increasing emphasis on environmental issues and fuel economy continue to drive turbine temperatures up. The higher engine operating temperatures results in an ever increasing severity of the operating environment inside a gas turbine. The severe operating environment results in more coating and base metal distress and increased maintenance costs. For example, more frequent replacement of the outer air seals.
A coating exists called a geometrically segmented abradable ceramic, (GSAC). The GSAC in development has the potential to satisfy the above described needs in many applications, however the current manufacturing methods are very costly with many added manufacturing steps including metallic layer buildup, diffusion heat treat and CNC milling of the divot structure. The complex surface geometry of certain parts and components creates a challenge to obtaining the best possible placement of surface features associated with the GSAC. There exists a need for a more effective process to locate the surface features of a GSAC on complex surface shapes.
In accordance with the present disclosure, there is provided a method of locating a surface feature for a geometrically segmented coating on a contoured surface comprising providing an article having a contoured surface; overlaying an assembly of ligaments on the contoured surface; and locating a surface feature at a predetermined location relative to each ligament.
In another embodiment, a method of locating a surface feature for a geometrically segmented coating on a contoured surface comprises providing an article having a contoured surface. The method including overlaying a triangular assembly of ligaments on the contoured surface and locating a surface feature at the center of each ligament.
In another embodiment, the method further comprises varying the size of the surface feature. The surface feature includes sharp corners and rounded corners. The article is a gas turbine engine component. The article is a gas turbine engine seal member.
In another embodiment, the assembly of ligaments is selected from the group consisting of a triangular assembly, a hexagonal assembly and a polygonal assembly.
In another embodiment, the predetermined location on each ligament is selected from the group consisting of a center of the ligament, and end of the ligament and a junction of ligaments.
In another alternative embodiment, a method of locating a surface feature for a geometrically segmented coating on a gas turbine engine component having contoured surfaces comprises providing the gas turbine engine component having at least one contoured surface; overlaying an assembly of ligaments on the contoured surface; locating a surface feature at a predetermined location relative to a portion of each said ligament; forming the surface feature in the contoured surface at the location; disposing a thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature.
In another embodiment, the method further comprises forming a pattern of features on the at least one contoured surface.
In another embodiment, the triangular assembly of ligaments conforms to the contoured surface. The surface feature comprises at least one of a sharp edge and a round edge. The surface feature comprises a sharp edge proximate a top and a bottom of the surface feature.
In another embodiment, the method further comprises disposing a bond coat layer to the at least one contoured surface, before disposing the thermally insulating topcoat.
In another embodiment, the gas turbine engine component is at least one of an airfoil, a seal, a bulkhead, a fuel nozzle guide, a transition duct and a combustor casing.
In another embodiment a method of locating a surface feature for geometrically segmented coatings on a gas turbine engine component having contoured surfaces comprises providing the gas turbine engine component having at least one contoured surface; overlaying an assembly of ligaments on the contoured surface; identifying a location for a surface feature on at least one of a geometric center of the assembly of ligaments and an intersection of the assembly of ligaments, a center of the ligaments, an end of the ligaments, and a combination of center and end of the ligaments; forming the surface feature in the contoured surface at the location; disposing a thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature.
In another embodiment the surface feature comprises at least one of a triangular shape, a circular shape, a hexagonal shape and a polygonal shape.
In another embodiment, the method further comprises forming the shaped surface feature from at least one of milling, laser engraving, casting, chemical etching and additive manufacturing.
In another embodiment, the method further comprises disposing a bond coat layer to the at least one contoured surface, before disposing the thermally insulating topcoat.
In another embodiment, the gas turbine engine component is at least one of an airfoil, a seal, a bulkhead, a fuel nozzle guide, a transition duct and a combustor casing.
In another embodiment, the assembly of ligaments is selected from the group consisting of a triangular assembly, a hexagonal assembly and a polygonal assembly.
Other details of the method are set forth in the following detailed description and the accompanying drawing wherein like reference numerals depict like elements.
Referring now to the
The seal member 30 includes a substrate 50, a plurality of divots, or geometric features or surface geometry 52 (hereinafter “features”) that are formed in a surface 54 on the gas path side of the seal member 30, and a thermally insulating topcoat 56 (e.g., a thermal barrier or TBC) disposed over the plurality of features 52 formed in the surface 54. The features 52 may not be shown to scale. The substrate 50 may include attachment features (not shown) for mounting the seal member 30 within the gas turbine engine 10.
The thermally insulating topcoat 56 includes segmented portions 60 that are separated by faults 62 extending through the thickness of the thermally insulating topcoat 56 from the features 52. The faults 62 extend from the edges or sides of the features 52 and facilitate reducing internal stresses within the thermally insulating topcoat 56 that may occur from sintering of the topcoat material at relatively high temperatures within the turbine section 20 during use in the gas turbine engine 10. Depending on the composition of the topcoat 56, surface temperatures of about 2500 degrees Fahrenheit (1370 degrees C.) and higher may cause sintering. The sintering may result in densification and diffusional shrinkage of the thermally insulating topcoat 56 and thereby induce internal stresses. In conventional non-segmented coatings the internal stresses due to sintering shrinkage may be high enough to cause spallation of the coating. In GSAC coating, the faults 62 provide pre-existing locations for accommodating the strain associated with sintering, reducing the internal stresses (e.g., reducing shear and radial stresses). That is, the energy associated with the internal stresses is maintained at a lower level due to the faults 62 such that there is less energy available for causing delamination cracking between the thermally insulating topcoat 56 and the underlying surface 54, substrate 50 or a bond coat 64 and spallation. The faults 62 facilitate reduction of internal stress energy within the thermally insulating topcoat 56.
The faults 62 may be produced by using any of a variety of different geometric surface features formed in the surface 54. In an exemplary embodiment, the surface features can comprise at least one of a triangular shape, a circular shape, a hexagonal shape, and any other polygonal shape. The triangular shape can include pointed or sharp ends and/or rounded ends/corners. In some manufacturing processes, such as, end mill or etching processes the corners can be rounded. The pattern and shape of the features 52 is not limited to any particular pattern and may be a grid type of pattern with individual perforations that extend from one surface of the sheet to the other surface of the surface 54. For example, a hexagonal close packed pattern of perforations may be formed in a sheet. Other patterns can include triangle, and polygonal patterns. The perforations being 0.080 inches in diameter and spaced on center at 0.105inch spacing.
The feature 52 forming process is selected to produce edges or sharp corners 58. Sharp corners at both the top and bottom of the GSAC divots are necessary for producing the necessary coating segmentation structure 60, 62. In an exemplary embodiment, the sharp corners 58 can be defined by the sum of the two radii less than or equal to 50 percent of the feature 53 height/depth. In another exemplary embodiment, the process can form sharp corners and/or rounded corners to any degree or combination as necessary to produce the coating segmentation structure 60.
The geometric surface features 52 may be selected to be any of a variety of different patterns or shapes. As an example, the features 52 may be formed as hexagonal walls that define a cell structure therebetween. Alternatively, the walls may be other shapes and need not be continuous.
The material selected for the substrate 50, bond coat 64 (if used), and thermally insulating topcoat 56 are not necessarily limited to any particular kind. For application on the seal member 30, the substrate 50 may be a metal alloy, such as a nickel based alloy.
The bond coat 64 may include any suitable type of bonding material for attaching the thermally insulating topcoat 56 to the surface 54. In some embodiments, the bond coat 64 includes a nickel alloy, platinum, gold, silver, or MCrAlY where the M includes at least one of nickel, cobalt, iron, or combination thereof, Cr is chromium, Al is aluminum and Y is yttrium. The bond coat 64 may be approximately 0.005 inches thick (approximately 0.127 millimeters), but may be thicker or thinner depending, for example, on the type of material selected and requirements of a particular application.
The thermally insulating topcoat 56 may be any type of ceramic material suited for providing a desired heat resistance in the gas turbine article. As an example, the thermally insulating topcoat 56 may be an abradable coating, such as yttria stabilized with zirconia, hafnia, and/or gadolinia, gadolinia zirconate, molybdate, alumina, or combinations thereof. The topcoats 56 may also include porosity. While various porosities may be selected, typical porosities in a seal application include 5 to 70% by volume.
Location of the surface features 52 may be optimized by way of a method that can be explained via a simple example, envision the structure of a geodesic dome 69 whose sub structure is a triangular assembly 70 of straight ligaments 72. (See
Varying feature diameter will help to maintain design criteria for the desired ratios of coating thickness to feature diameter, depth and inter-divot spacing.
In an alternative embodiment, the features can be located at locations in the center of the triangle formed by the intersection of the ligaments 72. The features 52 can be formed as triangle shaped features 84. This geometry would be best produced by casting, chemical etching or additive manufacturing. More complex shapes can of course be meshed and covered with features for formation of GSAC-TBC.
There has been provided a method of producing an array of surface features (surface geometry) that are adapted to complex surfaces by using a meshing method similar to those used to mesh a surface in finite element modeling. The method provides for an economical manufacturing process for GSAC part geometry. This method would be very applicable to use on large and simple geometry industrial gas turbine parts. The inventive method is a technique utilized to map out the best fit for the placement of the surface features that form the GSAC instead of using Cartesian methods to locate the surface features.
The use of the triangle geometry in the location method allows for a best fit of the surface features for the irregular shapes found on components, such as, airfoils and other engine parts. While the method has been described in the context of specific embodiments thereof, other unforeseen alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. The present invention provides the method and resultant product of geometrically segmented ceramic thermal barrier coating (GSC-TBC) on complex geometry.
The proven temperature/durability improvement achieved by this surface geometry method will now be available on complex shapes. This will improve the durability of air plasma sprayed TBC or permit operation at higher temperature or with lower cooling air flow rate. Accordingly, it is intended to embrace those alternatives, modifications, and variations which fall within the broad scope of the appended claims.
This application claims the benefit of provisional application Ser. No. 62/033,739, filed Aug. 6, 2014.
Number | Date | Country | |
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62033739 | Aug 2014 | US |