Patent Application
20150275682
1. Field of the Invention
This invention relates generally to providing a layer of alloy 230 on a base substrate of a component in a gas turbine engine and, more particularly, to providing a thin layer of alloy 230 sprayed on desired areas of the base substrate of components in a gas turbine engine before a bond coat layer and thermal barrier coating is applied to the substrate so as to reduce spallation of the thermal barrier coating.
2. Discussion of the Related Art
The world's energy needs continue to rise which provides a demand for reliable, affordable, efficient and environmentally-compatible power generation. A gas turbine engine is one known machine that provides efficient power, and often has application for an electric generator in a power plant, or engines in an aircraft or a ship. A typically gas turbine engine includes a compressor section, a combustion section and a turbine section. The compressor section provides a compressed air flow to the combustion section where the air is mixed with a fuel, such as natural gas, and ignited to create a hot working gas. The working gas expands through the turbine section where it is directed across rows of blades therein by associated vanes. As the working gas passes through the turbine section, it causes the blades to rotate, which in turn causes a shaft to rotate, thereby providing mechanical work.
Gas turbine engines of this type are periodically serviced for maintenance purposes. One of the maintenance operations is to detect erosion, mechanical fatigue and cracking in various turbine parts especially components in the hot gas path of the turbine section of the engine. The hot working gas paths for the first and second rows of blades, vanes and ring segments in the turbine section is directly from the combustion section of the engine, which frequently causes erosion and other damage of the these components at various locations and triggers thermal mechanical fatigue cracking. This causes the vanes to be reshaped, thus possibly directing the working gas in a non-optimal direction and could cause catastrophic failure.
A typical hot gas path component for a gas turbine engine has a base substrate that is a cast part made from a suitable nickel or cobalt alloy to withstand the high temperature environment of the turbine engine. Typically, the base substrate alloy, such as IN738, ECY-768 and IN939, includes a low concentration of aluminum so as to allow it to be easily welded when assembled and repaired when damaged. However, other vane design concerns would prefer that the base substrate alloy having a higher concentration of aluminum. For example, a thermal barrier coating (TBC), such as a suitable ceramic, is typically deposited on the base substrate to provide increased thermal protection for the vane. For those vanes employing a low aluminum alloy base substrate, a bond coat layer having a higher concentration of aluminum, such as a metallic coating layer, is generally deposited on the base substrate to provide the aluminum to help keep the TBC from spalling off of the vane. Particularly, as the base metal in these components is heated during operation of the gas turbine engine, the aluminum in the bond coat layer is oxidized creating alumina (Al2O3), which causes the aluminum in the bond coat layer to be depleted. Some of that alumina is formed at the transition between the bond coat layer and the TBC, which desirably operates to hold the TBC on the vane and preventing spallation of the TBC. However, a significant portion of that alumina is drawn into the substrate alloy because it has a low concentration of aluminum, which is thus not available to prevent the TBC from spalling off, and hence the TBC is lost over time. Therefore, as described, the lower the aluminum concentration in the base substrate alloy of the vane, the better the metal is for weldability and repairability, but the lower aluminum concentration increases spallation of the TBC.
The configuration of a low aluminum concentration base substrate, bond coat layer and TBC, especially for row 1 blades and row 1 and 2 vanes, often results in heavy oxidation leading to cracking in certain areas of the vane. This mainly results from early spallation of the TBC resulting in heavy oxidation, which is a result of overheating, and which depletes the bond coat of aluminum. When the TBC spalls, the metal temperature increases even more, which leads to burning and loss of metal. Further exasperating this failure is the fact that low aluminum vanes have poor oxidation resistances. When overheating occurs, a rapid loss of metal occurs, which causes a large loss of metal where the vanes mate together. Even further exasperating the problem is that vane alloys are notorious for their tendency to rapidly deplete the bond coat of aluminum. Compared to superior coating-compatible super alloys, such as CM 247, these vane alloys will always cause spallation of the TBC over relative shorter times. With the higher firing temperatures on the newer gas turbine designs, spallation could be seen on a CM247 alloy as well.
Known methods to address this problem include increasing the cooling air flow, which would decrease the metal temperature, but would also adversely impact engine power and efficiency in addition to increasing NOx emissions from the engine. Further, the vanes could be manufactured from a higher oxidation resistant alloy. However, this alternative has the drawback of increased cost, decreased castability and a decrease in weldability.
The present disclosure describes a technique for improving the thermal protection against oxidation for a component in a gas turbine engine, for example, blades, row 1 vanes and row 2 vanes. The technique includes spraying a thin layer of alloy 230 on a base substrate of the component at those locations on the component where thermal protection against oxidation is desired. A metal bond coat layer is then deposited on the alloy 230 layer and a thermal barrier coating is deposited on the bond coat layer. The chromium, molybdenum, iron and tungsten in alloy 230 provide superior oxidation resistance, and the addition of lanthanum in the alloy 230 helps tailor thermal expansion with the thermal barrier coating resulting in higher spallation life.
Additional features of the present invention will become apparent from the following description and appended claims, taken in conjunction with the accompanying drawings.
The following discussion of the embodiments of the invention directed to a technique for providing better thermal protection of a component in a gas turbine engine is merely exemplary in nature and is in no way intended to limit the invention or its applications or uses. For example, the technique described herein has particular application for blades and vanes in the engine. However, as will be appreciated by those skilled in the art, the technique may have application for other high temperature components in the engine.
Each group of the circumferentially disposed stationary vanes defines a row of the vanes and each group of the circumferentially disposed blades 34 defines a row 38 of the blades 34. In this non-limiting embodiment, the turbine section 16 includes four rows 38 of the rotating blades 34 and four rows of the stationary vanes in an alternating sequence. In other gas turbine engine designs, the turbine section 16 may include more or less rows of the turbine blades 34, It is noted that the most forward row of the turbine blades 34, referred to as the row 1blades, and the vanes, referred to as the row 1 vanes, receive the highest temperature of the working gas, where the temperature of the working gas decreases as it flows through the turbine section 16.
In this specific design, cooling holes 50 are provided in the air foils 46 and 48 and provide air cooling for those components. However, in this design, generally for increased power and performance requirements, there are no cooling holes provided in the inner shroud 42. Because the suction side 44 of the inner shroud 42 and the suction side of the outer shroud are subjected to very hot temperatures, and for those designs that employ a base substrate alloy that is low in aluminum, such as IN939, significant oxidation occurs at this area in the assembly 40 requiring significant repair, such as welding on replacement pieces, at periodic intervals, which is very costly.
The present invention proposes providing an oxidation resistant layer in addition to the known layers on the vane assembly 40, and other hot gas components, at the necessary and/or desired locations that is compatible with the known base alloy, bond coat layer and TBC layer. Particularly, the new layer is an alloy 230 layer, such as Haynes™ 230 alloy, deposited directly on the base substrate and before the bond coat layer. Alloy 230 is a known oxidation resistant material and has a chemical composition of 57 weight percent Ni, 22 weight percent Cr, 14 weight percent W, 2 weight percent Mo, 3 weight percent Fe, 5 weight percent Co, 0.5 weight percent Mn, 0.4 weight percent Si, 0.3 weight percent Al, 0.1 weight percent C, 0.02 weight percent La, and 0.015 weight percent B. The chromium, molybdenum, iron and tungsten in alloy 230 provide superior oxidation resistance, and the addition of lanthanum in the alloy 230 helps tailor thermal expansion with the thermal barrier coating resulting in higher spallation life.
It is noted that the discussion above refers to providing an alloy 230 layer at certain locations on the vane assembly 40 where greater thermal protection is desired. However, the alloy 230 layer can be applied to the entire vane assembly 40, including the airfoils 46 and 48, if thermal protection of those areas is desired. Further, the alloy 230 layer can be applied to other hot gas components including the blades 34 of the gas turbine engine 10, including the entire component, only certain blade portions or certain blade rows as necessary or desired.
The foregoing discussion discloses and describes merely exemplary embodiments of the present invention. One skilled in the art will readily recognize from such discussion, and from the accompanying drawings and claims, that various changes, modifications and variations can be made therein without departing from the scope of the invention as defined in the following claims.
