BACKGROUND OF THE INVENTION
This invention generally relates to protective coatings for metal alloy components exposed to high temperature gas environments and severe operating conditions, such as the working components of gas turbine engines used in electrical power generation. More particularly, the invention relates to a thermal barrier coating (TBC) for use in gas turbine engines and a method for producing a TBC coating.
The operating conditions to which gas turbine hardware components are exposed may be thermally and chemically severe. The surfaces of the metal substrates used to form turbine, combustor and augmentor components should exhibit greater than average mechanical strength, durability and erosion resistance in a very hostile, high temperature gas environment. “Erosion” generally refers to the process whereby a surface, particularly metal, is bombarded by contaminant particles of sufficiently high energy that cause other particles to be ejected (eroded) from the surface, resulting in degradation and cracking of the substrate material.
Recent advances have been achieved by using high temperature alloys in gas turbine systems by incorporating iron, nickel and cobalt-based superalloys in coatings applied to the substrate of key turbine components. The purpose of an effective surface coating is generally two-fold. First, the coating should form a protective and adherent layer that guards the underlying base material against oxidation, corrosion, and degradation. Second, the coating should have low thermoconductivity relative to the substrate. As superalloy compositions have become more complex, it has been increasingly difficult to obtain both the higher strength levels that are required (particularly at increased gas turbine operating temperatures) and a satisfactory level of corrosion and oxidation resistance. The trend towards higher gas turbine firing temperatures has made the oxidation, corrosion and degradation problems even more difficult. Thus, despite recent improvements in thermal barrier coatings, a significant need may still exist for more cost-effective, more efficacious, and less degradable high temperature coatings, because many alloy components cannot withstand the long service exposures and repetitive cycles encountered in a typical gas turbine environment.
Many of the known prior art coatings used for gas turbine components include aluminide and ceramic components. Typically, ceramic coatings have been used in conjunction with a bond coating formed from an oxidation-resistant alloy such as MCrAlY, where M is iron, cobalt, and/or nickel, or from a diffusion aluminide or platinum aluminide that forms an oxidation-resistant intermetallic. In higher temperature applications, these bond coatings form an oxide layer or “scale” that chemically bonds to the ceramic layer to form the final bond coating.
It has also been known to use zirconia (ZrO2) that is partially or fully stabilized by yttria (Y2O3), magnesia (MgO) or other oxides as the primary constituent of the ceramic layer. Yttria-stabilized zirconia (YSZ) is often used as the ceramic layer for thermal bond coatings because it may exhibit favorable thermal cycle fatigue properties. That is, as the temperature increases or decreases during gas turbine start up and shut down, the YSZ is capable of resisting stresses and fatigue much better than other known coatings. Typically, the YSZ is deposited on the metal substrate using known methods, such as air plasma spraying (APS), low pressure plasma spraying (LPPS), as well as by physical vapor deposition (PVD) techniques such as electron beam physical vapor deposition (EBPVD). Notably, YSZ deposited by EBPVD is characterized by a strain-tolerant columnar grain structure that enables the substrate to expand and contract without causing damaging stresses that lead to spallation. The strain-tolerant nature of such systems may be known. See generally U.S. Pat. No. 6,730,413 for a description of a known thermal barrier coating system.
The production of vertical cracks in a manufacturing environment may be difficult and/or problematic. In certain aspects, the present invention may generally relate to a process involving inducing cracks or microcracks in a post-coating application. This may facilitate the thermal barrier coating be applied densely, which may be easier to accomplish. After application, the coating may be selectively cracked, e.g., using shockwave exposure.
Laser peening is well known and understood in the art. For example, laser peening has been used to create a compressively stressed protection layer at the outer surface of a workpiece which is known to considerably increase the resistance of the workpiece to fatigue failure as disclosed in U.S. Pat. No. 4,937,421. Laser shock peening has also been used create deep compressive residual stresses into a turbine blade as disclosed in U.S. Pat. No. 5,591,009.
BRIEF DESCRIPTION OF THE INVENTION
In an aspect, an embodiment may generally relate to a method for forming cracks in a thermal barrier coating applied to a gas turbine component. The method may include the following steps: depositing a bond coating on a metallic substrate, wherein the bond coating comprises MCrAlY, where M is iron, cobalt, and/or nickel, and wherein the metallic substrate comprises a gas turbine component; depositing a thermal barrier coating on the bond coating, wherein the thermal barrier coating comprises yttria-stabilized zirconia; subjecting at least a portion of the thermal barrier coating to a shockwave such that microcracks are formed in the thermal barrier coating and such that the metallic substrate is not substantially deformed.
In an aspect, an embodiment may generally relate to a method of forming cracks in a ceramic-based coating. The method may include the following steps: depositing a ceramic-based coating on a metallic-based substrate, wherein the ceramic-based coating comprises a thermal barrier coating; and subjecting at least a portion of the ceramic-based coating to a shockwave such that microcracks are formed in the ceramic-based coating and such that the metallic-based substrate is not substantially deformed.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a cross-sectional view of a metal substrate, such as a high pressure gas turbine blade, showing the thermal barrier coating as applied to the blade using a laser shock process in accordance with an embodiment of the invention.
FIG. 2 schematically illustrates the amount of energy required to induce cracks in a thermal barrier coating.
DETAILED DESCRIPTION OF THE INVENTION
As noted above, thermal barrier coatings according to the present invention are applicable to various metal alloy components (so-called “superalloys”) that must still be protected from a thermally and chemically hostile environment. Examples of such components include nozzles, buckets, shrouds, airfoils, and other hardware found in almost any gas turbine engine.
The coating may be any known TBC composition, e.g., it may consist of a thermal insulating ceramic layer whose composition and deposition significantly enhance the erosion resistance of the turbine components while maintaining a spallation resistance equivalent to or better than conventional coatings. The coating composition may be applied then cracked after application.
High pressure turbine blades are prime examples of the substrates to which coatings in accordance with the invention can be applied. Typically, turbine blades have an airfoil and a platform against which hot combustion gases are directed during operation of the gas turbine. Thus the airfoil surfaces are subjected to attack by oxidation, corrosion, and erosion. The airfoil normally is anchored to a turbine disk with a dovetail formed on a root section of the blade.
FIG. 1 shows a thermal barrier coating in accordance with the invention as applied to a substrate. The coating 10 includes a thermal-insulating ceramic layer 12 over a bond coating 14 that overlies a metal alloy substrate 16 which may form the base material of the turbine blade. Suitable materials for the substrate include iron-, nickel-, and/or cobalt-based superalloys. The bond coating may be oxidation resistant and may form an alumina layer 18 on the surface of the bond coating when the coated blade is exposed to elevated temperatures. The alumina layer may protect the underlying superalloy substrate 16 from oxidation and may provide a surface to which the ceramic layer adheres.
Within layer 12, there are vertical cracks that have been formed so as to increase and/or induce strain tolerance. Crack induction via shockwave exposure may enable the cracks to be placed in the material in particularly desirable areas and at specifically desirable densities. To form the cracks, coupled ablation may be used to induce a shockwave into a material. The coupled ablation may be achieved through the use of a pulsed laser in a process similar to laser shock peening, where a laser is pulsed thorough the coupling material and into the ablative material thus creating a shockwave.
In the prior art, laser shock peening may be used to densify the material. In the case of a TBC, though, the resultant shockwave can induce microcracks within the coating to provide strain tolerance. Other means of shockwave exposure may be possible. Other means of coupled ablation may also be possible.
In an exemplary embodiment, a strain tolerant TBC may be formed using laser shock peening. A thermal barrier coating may be applied to a metallic substrate using an air plasma spray. A bond coat may be MCrAlY (where M is iron, cobalt, and/or nickel), and the TBC may be 8% yttria-stabilized zirconia or any other ceramic-based coating used as a thermal barrier on turbine components. After application to the substrate, the TBC may be laser shock peened.
The energy used to induce the microcracks in the TBC should preferably not substantially deform the substrate. Thus, the energy should be relatively low because the coating may be very thin. In order to substantially deform the substrate, the energy of the shock wave would need be to sufficient to impart stress at or above the plastic yield of the substrate but below its compressive strength. In contrast thereto, the energy used to induce microcracks in the TBC should be sufficient to impart stress above compressive strength of the TBC. Because the metallic substrate may be ductile, and the ceramic TBC may be brittle, there may be a particular level of energy that can be selected or determined.
FIG. 2 schematically illustrates a general description of the amount of energy required to induce cracks in a TBC. The amount of energy (per unit area) to fracture a material is represented by the area under the stress/strain curve. FIG. 2 illustrates a typical porous TBC coating. The porosity reduces the “effective” cross-sectional area and therefore reducing the force required for fracture (because energy is a function of force not pressure or stress). This may effectively reduce the area under the curve considerably.
In preferred embodiments, a thermal barrier coating experiences a shockwave (e.g., via laser ablation) and is fractured. The energy that may be required may depend on the source of the shockwave, e.g., laser ablation or other, and/or the properties of material being cracked.
Thus, in certain embodiments, a process (e.g., laser ablation or laser shock peening) may produce a microstructural features (e.g., vertical cracks). This may increase the durability of a turbine component and/or reduce manufacturing costs. For example, a simple dense coating may be applied to a component, and the vertical cracks can be induced in areas that they are needed. That is, cracks need not be introduced throughout an entire coating via processing parameters.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.