This invention relates generally to ceramic coating layers and more particularly, to dense, vertically cracked thermal barrier coating layers.
In some known gas turbine engines, some components such as bucket airfoils, may be subjected to high temperature conditions in excess of 1000 degrees Celsius (° C.) (1832 degrees Fahrenheit (° F.)) when in service. In order to protect these components, a ceramic thermal barrier coating (TBC) layer may be used to provide an effective and reliable thermally insulating barrier between the base metal, or ceramic, substrate of the components and high temperature environments. As is known in the art, the smoothness of the coating may affect the aerodynamic properties of the surface as well as facilitate reducing heat transfer coefficient.
In some known components, a TBC layer is formed by a plasma spray process to achieve desired structural characteristics, i.e., mechanical and thermal properties. A typical plasma spray process may involve use of a plasma spray torch, or nozzle, that produces hot ionized plasma for melting TBC powder injected therein.
One known TBC layer deposition is often referred to as a dense, vertically cracked (DVC) material. DVC TBC materials tend to have a dense microstructure having an accompanying low porosity that facilitates increased erosion resistance. DVC TBC materials also have a plurality of vertical micro-cracks that facilitate increased strain tolerance. However, the high densities and low porosities may tend to increase the difficulty associated with smoothing the surfaces of components that have received a DVC TBC layer.
Some known component manufacturing processes involve use of some known “smooth coat” ceramic TBC materials to form a layer on the components subsequent to the DVC TBC application. At least some of such smooth coat materials produce smooth surfaces on porous thermal barrier coatings. However, some of these smooth coat materials typically may not adhere reliably to DVC TBC. As a result, the smooth coat layer may “spall” (i.e., delaminate, or lift off) from the component surface during curing, thereby damaging the component prior to completion of manufacturing.
In one aspect, a method for manufacturing a machine component is provided. The method includes forming a machine component substrate wherein the substrate has a substrate surface region. The method further includes forming at least one primary thermal barrier layer wherein the at least one primary thermal barrier layer includes a first primary thermal barrier layer. The first primary thermal barrier layer further includes a first ceramic thermal barrier material having a first porosity. The method also includes forming at least one secondary thermal barrier layer wherein the at least one secondary thermal barrier layer is formed over at least a portion of the first primary thermal barrier layer. Also, the secondary thermal barrier layer further includes a second ceramic thermal barrier material having a second porosity, wherein the second porosity is greater than the first porosity. The method also includes forming at least one tertiary thermal barrier layer comprising a smooth coat material having a tertiary porosity, wherein the tertiary thermal barrier layer is formed over at least a portion of the secondary thermal barrier layer. The secondary thermal barrier layer facilitates reducing a delamination of the tertiary thermal barrier layer. The method further includes curing the tertiary thermal barrier layer.
In another aspect, a method for manufacturing a turbine component having a thermal barrier coating is provided. The method includes forming a turbine component substrate wherein the substrate has a substrate surface region. The method further includes forming at least one primary thermal barrier layer using a spray nozzle positioned a first distance from the component. The forming of the primary thermal barrier layers include the formation of a first primary thermal barrier layer over at least a portion of the substrate surface region of the machine component. At least one subsequent primary thermal barrier layer extends substantially over each previously formed primary thermal barrier layer. The plurality of primary thermal barrier layers include a first ceramic thermal barrier material having a first porosity. The method also includes forming at least one secondary thermal barrier layer using a spray nozzle positioned a second distance from the component wherein the second distance is greater than the first distance. The secondary thermal barrier layer is formed over at least a portion of the primary thermal barrier layers. Furthermore, the secondary thermal barrier layer includes a second ceramic thermal barrier material having a second porosity that is greater than the first porosity. The method also includes forming at least one tertiary thermal barrier layer that includes a smooth coat material having a tertiary porosity. The tertiary thermal barrier layer is formed over at least a portion of the at least one secondary thermal barrier layer. The method further includes curing the tertiary thermal barrier layer in air at a predetermined temperature wherein the secondary thermal barrier layer facilitates reducing a delamination of the tertiary thermal barrier layer.
In a further aspect, a machine component is provided. The machine component includes a substrate comprised of a surface region wherein the substrate further includes an article having predetermined dimensions. The component also includes at least one primary thermal barrier layer that has a first porosity. The component further includes at least one secondary thermal barrier layer that has a second porosity, wherein the second porosity is greater than the first porosity. The component also includes at least one tertiary thermal barrier layer having a tertiary porosity, wherein the second porosity facilitates reducing a delamination of the at least one tertiary thermal barrier layer.
As used herein, the term layer refers to, but is not limited to, a sheet-like expanse or region of a material or materials covering a surface, or forming an overlying or underlying part or segment of an article such as a turbine component. A layer has a thickness dimension. The term layer does not refer to any particular process by which the layer is formed. For example, a layer can be formed by spraying, coating, or a laminating process.
Substrate 102 includes surface region 104 and may be shaped with predetermined dimensions to a set of predetermined contours and thicknesses substantially similar to the dimensions of finished turbine component 100. In the exemplary embodiment, substrate 102 may be metallic. Alternatively, substrate 102 may be ceramic.
In the exemplary embodiment DVC TBC layers 105 includes eight (8) layers with first DVC TBC layer 106 formed over surface region 104, and seven (7) subsequent DVC TBC layers 107, each subsequent layer formed over a previous layer. The thickness and porosity of each layer are substantially similar. The thickness of each of plurality of layers 105 is approximately 0.0508 millimeters (mm) (0.002 inches) each for a total thickness of approximately 0.4064 mm (0.016 inches). Circumferentially outermost surface 109 of layer 108 may be smoothed. DVC TBC layers 105 may be a metal oxide material, such as yttria-stabilized zirconia having a chemical composition of 6-8 weight percent yttria with a balance of zirconia. Alternately, DVC TBC layers 105 may include other ceramic materials and the associated number of layers and the thicknesses of these layers may be varied according to appropriate standards and tolerances. In the exemplary embodiment, one secondary TBC layer 110 is formed over circumferentially outermost DVC TBC layer surface 108. Layer 110 is a layer that is less dense, i.e., more porous, than DVC TBC layers 105.
In the exemplary embodiment, one tertiary TBC layer 114 is formed over surface 112. Layer 114 is approximately 0.0254 mm (0.001 inches) to 0.0508 mm (0.002 inches). In the exemplary embodiment a smooth coat material is used, such as, but not limited to, AJ11. In the exemplary embodiment, one layer with the thickness described above is formed with a surface roughness that may be approximately less than 2.54 micrometers (100 micro-inches) roughness average (RA). Alternatively, the number of layers and the thickness of the layers may be varied according to the component's 100 operational application.
Method 200 includes a method step 204 that further includes forming at least one primary DVC TBC layer 105 (shown in
In the exemplary embodiment DVC TBC layers 105 may be deposited with eight (8) spray passes with the torch or nozzle located a distance of approximately 11.43 centimeters (cm) (4.5 inches) from substrate surface 104, using a computer-controlled program with robotic motion for reproducibility. The thickness and porosity of each of layers 105 are substantially similar. This process produces a uniformly hard, dense, ceramic coating, adding about 0.0508 millimeters (mm) (0.002 inches) per pass for a total thickness of approximately 0.4064 mm (0.016 inches). In the exemplary embodiment, surface 109 is not smoothed. In an alternative embodiment, the aforementioned total thickness allows for approximately 0.0508 mm (0.002 inches) to be removed during finishing operations of layer 108 (shown in
The ceramic material used to form plurality of DVC TBC layers 105 may be a metal oxide, such as yttria-stabilized zirconia having a chemical composition of 6-8 weight percent yttria with a balance of zirconia. Alternately, other ceramic materials may also be used.
In the exemplary embodiment, method 200 includes a method step 206 that further includes forming one secondary TBC layer 110 (shown in
A method step 210 of exemplary method 200 includes applying a high temperature heat treatment to component 100. The associated heat treatment temperatures and time periods may vary based on a plurality of parameters that may include, but not be limited to, the number of layers and the predetermined thicknesses of the layers.
A method step 212 of exemplary method 200 includes forming a tertiary TBC layer 114 (shown in
A method step 214 of exemplary method 200 includes heat curing component 100 in air at a temperature of approximately 900° C. (1650° F.). This step is a test in that heating the component induces stresses that may delaminate tertiary layer 114 from secondary layer 110 if the adherence of layer 114 to layer 110 is not sufficient. The associated heat curing temperatures and time periods may vary based on a plurality of parameters that may include, but not be limited to, the number of layers and the predetermined thicknesses of the layers.
The component manufacturing methods described herein facilitate application of a protective thermal barrier layer to a component. More specifically, forming a plurality of protective layers on the turbine component described above prevents damage in high temperature environments. As a result, degradation of the component when placed in service and increased manufacturing costs may be reduced or eliminated.
Although the methods described and/or illustrated herein are described and/or illustrated with respect to manufacturing a component, and more specifically, a turbine component, practice of the methods described and/or illustrated herein is not limited to turbine components nor to forming thermal barrier layers generally. Rather, the methods described and/or illustrated herein are applicable to manufacturing any article and forming any layer of any material.
Exemplary embodiments of turbine component manufacturing are described above in detail. The methods, apparatus and systems are not limited to the specific embodiments described herein nor to the specific turbine components manufactured, but rather, the methods of manufacturing turbine components may be utilized independently and separately from other methods, apparatus and systems described herein or to manufacturing components not described herein. For example, other components can also be manufactured using the methods described herein.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.