The present invention relates to components designed to operate at high temperatures. More particularly, this invention relates to airfoils for gas turbine engines and methods for manufacture and repair of such components.
In a gas turbine engine, compressed air is mixed with fuel in a combustor and ignited, generating a flow of hot combustion gases through one or more turbine stages that extract energy from the gas, producing output power. Each turbine stage includes a stator nozzle having vanes which direct the combustion gases against a corresponding row of turbine blades extending radially outwardly from a supporting rotor disk. The vanes and blades include airfoils having a generally concave xe2x80x9cpressurexe2x80x9d side and a generally convex xe2x80x9csuctionxe2x80x9d side, both sides extending axially between leading and trailing edges over which the combustion gases flow during operation. The vanes and blades are subject to substantial heat load, and, because the efficiency of a gas turbine engine is proportional to gas temperature, the continuous demand for efficiency translates to a demand for airfoils that are capable of withstanding higher temperatures for longer service times.
Gas turbine airfoils on such components as vanes and blades are usually made of superalloys and are often cooled by means of internal cooling chambers and the addition of coatings, including thermal barrier coatings (TBC""s) and environmentally resistant coatings, to their external surfaces. The term xe2x80x9csuperalloyxe2x80x9d is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more other elements including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The internal air cooling of turbine airfoils is often accomplished via a complex cooling scheme in which cooling air flows through channels within the airfoil (xe2x80x9cinternal air cooling channelsxe2x80x9d) and is then discharged through a configuration of cooling holes at the airfoil surface. Convection cooling occurs within the airfoil from heat transfer to the cooling air as it flows through the cooling channels. In addition, fine internal orifices are often provided to direct cooling air flow directly against inner surfaces of the airfoil to achieve what is referred to as impingement cooling, while film cooling is often accomplished at the airfoil surface by configuring the cooling holes to discharge the cooling air flow across the airfoil surface so that the surface is protected from direct contact with the surrounding hot gases within the engine. TBC""s comprise at least a layer of thermally insulating ceramic and often include one or more layers of metal-based, oxidation-resistant materials (xe2x80x9cenvironmentally resistant coatingsxe2x80x9d) underlying the insulating ceramic for enhanced protection of the airfoil. Environmentally resistant coatings are also frequently used without a TBC topcoat. Technologies such as coatings and internal air cooling have effectively enhanced the performance of turbine airfoils, but material degradation problems persist in turbine airfoils due to locally aggressive conditions in areas such as airfoil leading edges and trailing edges.
A considerable amount of cooling air is often required to sufficiently lower the surface temperature of an airfoil. However, the casting process and the cores required to form the cooling channels limit the complexity of the cooling scheme that can be formed within an airfoil at leading and trailing edges of vanes and blades. The resulting restrictions in cooling airflow often promote higher local temperatures in these areas relative to those existing in other locations on a given airfoil, giving rise to increased oxidation in these areas. In typical jet engines, for example, bulk average airfoil temperatures range between about 898xc2x0 C. (about 1650xc2x0 F.) to about 982xc2x0 C. (about 1800xc2x0 F.), while airfoil leading and trailing edges often reach about 1149xc2x0 C. (about 2100xc2x0 F.) or more. Maximum temperatures are expected in future applications to be over about 1315xc2x0 C. (about 2400xc2x0 F.). At such elevated temperatures, the oxidation process consumes metal parts, forming a weak, brittle metal oxide that is prone to chip or spall away from the part. In addition to oxidation, erosion due to impact of particles entrained in the gas flow often occurs at airfoil leading edges, completely removing the TBC from the area and exposing the base metal to the hot gas flow, increasing the local rate of material removal due to oxidation and wear.
Erosion and oxidation of material at the edges of airfoils lead to degradation of turbine efficiency. As airfoils are worn away, gaps between components become excessively wide, allowing gas to leak through the turbine stages without the flow of the gas being converted into mechanical energy. When efficiency drops below specified levels, the turbine must be removed from service for overhaul and refurbishment. A significant portion of this refurbishment process is directed at the repair of the specific areas of the airfoils described above. For example, damaged material is removed and then new material built onto the blade by welding with filler material or by laser deposition of metal powders.
In current practice, the original edge material is made of the same material as the rest of the original blade, often a superalloy based on nickel or cobalt. Because this material was selected to balance the design requirements of the entire blade, it is generally not optimized to meet the special local requirements demanded by conditions at the airfoil leading or trailing edges. The performance of alloys commonly used for repair is comparable or inferior to that of the material of the original component, depending upon the microstructure of the repaired material, its defect density due to processing, and its chemistry. For example, many turbine airfoils are made using alloys that have been directionally solidified. The directional solidification process manipulates the orientation of metal crystals, or grains, as the alloy is solidified from the molten state, lining the grains up in one selected primary direction. The resultant alloy has enhanced resistance to high temperature deformation, referred to as xe2x80x9ccreep,xe2x80x9d during service when compared to conventionally processed materials. Advanced applications employ alloys made of a single crystal for even further improvements high temperature properties. However, when these components are repaired by conventional processes, using build-up of weld filler material, the resulting microstructure of the repair is typical of welded material, not directionally solidified or single-crystalline. Other repair methods, such as applying powder mixtures wherein one powder melts and densifies the repaired area during heat treatment, results in microstructures different from the parent alloy. Such microstructures, present in a conventional airfoil material such as a superalloy, may cause the airfoil to require excessively frequent repairs in advanced designs that rely on the benefits of directional solidification or single crystal processing to maintain performance.
The term xe2x80x9coxidation resistancexe2x80x9d is used in the art to refer to the amount of damage sustained by a material when exposed to oxidizing environments, such as, for example, high temperature gases containing oxygen. Oxidation resistance is generally measured as the rate at which the weight of a specimen changes per unit surface area during exposure at a given temperature. In many cases, the weight change is measured to be a net loss in weight, as metal is converted to oxide that later detaches and falls away from the surface. In other cases, a specimen may gain weight if the oxide tends to adhere to the specimen, or if the oxide forms within the specimen, underneath the surface, a condition called xe2x80x9cinternal oxidation.xe2x80x9d A material is said to have xe2x80x9chigherxe2x80x9d or xe2x80x9cgreaterxe2x80x9d oxidation resistance than another if the material""s rate of weight change per unit surface area is closer to zero than that of the other material for exposure to the same environment and temperature.
The so-called xe2x80x9cplatinum groupxe2x80x9d of metal elements comprises rhodium (Rh), osmium (Os), platinum (Pt), iridium (Ir), ruthenium (Ru), palladium (Pd), and rhenium (Re)elements noted for high chemical resistance. Several elements from this group are noteworthy as examples of materials with substantially higher oxidation resistance relative to current airfoil materials. Some platinum group metals and several alloys based on platinum group metals possess excellent resistance to oxidation at temperatures exceeding the capabilities of many Ni-based superalloys. The class of materials referred to as xe2x80x9crefractory superalloysxe2x80x9d offer additional strength over the platinum group metals, though at the expense of some oxidation resistance. These alloys are based on Ir or Rh, with transition metal additions of up to about 20 atomic percent, and are strengthened by a precipitate phase of generic formula M3X, where M is Rh or Ir and X is typically Ti, V, Ta, or Zr, or combinations thereof. Some alloys of this type can withstand 1-2 hour exposures to at least about 1593xc2x0 C. (about 2900xc2x0 F.) without catastrophic oxidation. Use of materials incorporating platinum-group metals has been restrained to date due to the high density and very high cost of these materials in comparison to more conventional airfoil materials.
The selection of a particular alloy for use in a given airfoil design is accomplished based on the critical design requirements for a number of material properties, including strength, toughness, environmental resistance, weight, cost, and others. When one alloy is used to construct the entire airfoil, compromises must be made in the performance of the airfoil because no single alloy possesses ideal values for the long list of properties required for the airfoil application, and because conditions of temperature, stress, impingement of foreign matter, and other factors are not uniform over the entire airfoil surface.
It would be advantageous if the performance of both newly manufactured and repaired airfoils could be improved to better withstand the aggressive environments present in localized areas on turbine components. However, it would not be desirable if improvements to environmental resistance were effected at the expense of other design critical requirements of the airfoil. For example, a blade made of pure platinum would have excellent oxidation resistance, but would lack the needed strength and would cost many times the price of a blade made of conventional superalloy material. Therefore, it would be beneficial if turbine airfoils could be improved in a manner that would allow for enhanced performance in regions susceptible to damage due to locally aggressive conditions without significantly detracting from the overall performance of the airfoil.
In one aspect of the present invention, a gas turbine airfoil is provided which comprises a wall, the wall defining the perimeter of the airfoil and comprising a leading edge section and a trailing edge section, wherein a majority of the surface area of the wall comprises a first material, the first material having an oxidation resistance and a melting temperature, and at least one portion of the wall comprises a second material, the second material having an oxidation resistance that is greater than the oxidation resistance of the first material and a melting temperature that is at least about 83 degrees Celsius (about 150 degrees Fahrenheit) greater than the melting temperature of the first material, the at least one portion of the wall located in at least one section of the wall selected from the group consisting of the leading edge section and the trailing edge section.
Another aspect of the invention provides a method for repairing a gas turbine airfoil, the method comprising: a. providing an airfoil comprising a wall, the wall defining the perimeter of the airfoil and comprising a leading edge section and a trailing edge section, and further comprising a first material with a melting temperature and an oxidation resistance; b. removing at least one portion of the wall, the at least one portion located in at least one section of the wall selected from the leading edge and the trailing edge; c. providing a second material, the second material having an oxidation resistance that is greater than the oxidation resistance of the first material and a melting temperature that is at least about 83xc2x0 C. (about 150 degrees Fahrenheit) greater than the melting temperature of the first material; and d. disposing the second material onto the wall in the at least one section where the at least one portion of the wall was removed.
Another aspect of the invention provides a method for manufacturing a gas turbine airfoil, the airfoil comprising a wall, the wall having a cross-sectional thickness that is specified to a nominal dimension, the wall defining the perimeter of the airfoil and comprising a leading edge section and a trailing edge section, and further comprising a first material with a melting temperature and an oxidation resistance, the method comprising: a. providing an airfoil with a deficit in cross-sectional wall thickness, relative to the specified nominal cross-sectional wall thickness dimension, in at least one section of the wall, the section selected from the leading edge section and the trailing edge section; b. providing a second material, the second material having an oxidation resistance that is greater than the oxidation resistance of the first material and a melting temperature that is at least about 83xc2x0 C. (about 150 degrees Fahrenheit) greater than the melting temperature of the first material; and c. disposing the second material onto the wall at the at least one section such that the deficit in cross-sectional wall thickness is eliminated.
Another aspect of the invention provides an insert for repair and manufacture of a gas turbine airfoil, the airfoil comprising a wall, the wall having an outer surface, the wall defining the perimeter of the airfoil and comprising a leading edge section and a trailing edge section, the wall further comprising a first material with a melting temperature and an oxidation resistance, the insert comprising an outer surface that is shaped such that the outer surface of the insert conforms with the outer surface of the wall at a section of the wall selected from the group consisting of the leading edge section and the trailing edge section, the insert comprising a second material, the second material having oxidation resistance that is greater than the oxidation resistance of the first material and a melting temperature that is at least about 83xc2x0 C. (about 150 degrees Fahrenheit) greater than the melting temperature of the first material.
USPTO Backgrounds
USPTO Backgrounds