Repair of gas turbine blade tip without recoating the repaired blade tip

This application is a divisional of U.S. application Ser. No. 11/022,185, filed Dec. 23, 2004, which is incorporated by reference in its entirety.

This invention relates to the repair of the damaged blade tip of a damaged gas turbine blade and, more particularly, to such a repair wherein the repaired blade tip is not recoated with a non-ceramic environmental coating or bond coat.

In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is combusted, and the resulting hot combustion gases are passed through a turbine mounted on the same shaft. The turbine includes a rotor body with a series of turbine blades extending radially outward from the rotor body, and a stationary shroud that forms a tunnel in which the rotor body and its blades turn. The flow of combustion gas turns the turbine by contacting an airfoil portion of the turbine blade, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. There may additionally be a bypass fan that forces air around the center core of the engine, driven by a shaft extending from the turbine section.

The turbine blades are currently made of nickel-base superalloys that have acceptable mechanical properties in the operating conditions of the gas turbine engine. Those nickel-base superalloys are usually coated with a protective coating that protects against oxidation damage. The protective coating includes a non-ceramic coating on the lateral surface of the airfoil. The protective coating may also include a ceramic layer that overlies the non-ceramic coating and insulates the turbine blade to allow it to function for longer times at higher temperatures than would otherwise be possible.

During service and despite the presence of the coating, the tips of some of the turbine blades may be damaged by rubbing contact with the stationary shroud of the gas turbine, by oxidation by the hot combustion gases, and by particle impacts. If the damage becomes sufficiently severe to a turbine-blade tip so that the dimensions of the turbine blade are reduced to less than the specified minimum values and/or the overall engine performance becomes unacceptable, the damaged turbine blade is removed from service. The damaged turbine blade may then be repaired and returned to service or discarded, but repair is preferred because of the high cost of each new turbine blade. The decision to repair or discard is in part economic, so that the higher the cost of the repair, the less likely that the turbine blade will be repaired and the more likely that an expensive new turbine blade will be installed.

In the conventional repair process as now practiced, adjacent protective coatings are removed, the damaged material at the tip is removed, repair material is applied to restore the dimensions of the turbine blade to the specified range, the lateral surface of the tip area is recoated, and the repaired-and-recoated turbine blade is heat treated. To improve overall gas-turbine-engine economics, there is a need to reduce the cost of the repair. The present invention fulfills this need, and further provides related advantages.

The present approach provides a method for repairing a gas turbine blade having a damaged blade tip region, and a repaired turbine blade. The present approach reduces the cost of the repair, by reducing the need for recoating of the repaired blade tip region.

A method for repairing a damaged gas turbine blade includes furnishing the damaged gas turbine blade which has previously been in service and which is made of a base metal. Any damaged material is removed from the damaged blade tip of the gas turbine blade. The damaged blade tip is weld repaired with a nickel-base superalloy repair alloy that is different from the base metal and is more resistant to oxidation than is the base metal in the operating environment of the gas turbine blade, to form a tip-repaired gas turbine blade having a repaired blade tip. The method does not include any step of coating a lateral surface of the repaired blade tip with a non-ceramic protective coating after the step of weld repairing. The method preferably does not include any step of coating the lateral surface of the repaired blade tip with a ceramic coating after the step of weld repairing, but optionally a ceramic coating may be applied.

The method optionally includes an additional step, after the step of weld repairing, of heat treating the tip-repaired gas turbine blade. If the heat treatment is performed, the tip-repaired gas turbine blade is preferably heat treated at a temperature of from about 1850° F. to about 2050° F. and for a time of from about 1 hour to about 8 hours, followed by an additional ageing heat treatment at a temperature of from about 1500° F. to about 1700° F. and for a time of from about 2 hours to about 16 hours. If a ceramic thermal barrier coating is applied, this same heat treatment may optionally be employed after the deposition of the ceramic coating.

A most preferred nickel-base superalloy used as the repair alloy to weld repair the damaged blade tip has a nominal composition, in weight percent, of from about 7.4 to about 7.8 percent chromium, from about 5.3 to about 5.6 percent tantalum, from about 2.9 to about 3.3 percent cobalt, from about 7.6 to about 8.0 percent aluminum, from about 0.12 to about 0.18 percent hafnium, from about 0.5 to about 0.6 percent silicon, from about 3.7 to about 4.0 percent tungsten, from about 1.5 to about 1.8 percent rhenium, from about 0.01 to about 0.03 percent carbon, from about 0.01 to about 0.02 percent boron, balance nickel and minor elements. Even more preferably, the nickel-base superalloy used to weld repair the damaged blade tip has a nominal composition in weight percent of about 0.01-0.03 percent carbon, 0.1 percent maximum manganese, about 0.5-0.6 percent silicon, 0.01 percent maximum phosphorus, 0.004 percent maximum sulfur, about 7.4-7.8 percent chromium, about 2.9-3.3 percent cobalt, 0.10 percent maximum molybdenum, about 3.7-4.0 percent tungsten, about 5.3-5.6 percent tantalum, 0.02 percent maximum titanium, about 7.6-8.0 percent aluminum, about 1.5-1.8 percent rhenium, 0.005 percent maximum selenium, 0.3 percent maximum platinum, 0.01-0.02 percent boron, 0.03 percent maximum zirconium, about 0.12-0.18 percent hafnium, 0.1 percent maximum niobium, 0.1 percent maximum vanadium, 0.1 percent maximum copper, 0.2 percent maximum iron, 0.0035 percent maximum magnesium, 0.01 percent maximum oxygen, 0.01 percent maximum nitrogen, balance nickel with other elements total 0.5 percent maximum.

The repair alloy preferably has an oxidation resistance equal to or better than these specific repair alloys of the preceding paragraph. The repair alloy may be a variation of one disclosed herein, such as one containing from about 1 to about 1.5 weight percent rhenium or from about 0.2 to about 0.6 weight percent hafnium. The repair alloy may instead be a completely different nickel-base superalloy.

The repaired gas turbine blade comprises a body including an airfoil made of a base metal, and a blade tip of the airfoil made of a nickel-base superalloy repair alloy of a different composition than the base metal. The repair alloy of the repaired blade tip is more resistant to oxidation than is the base metal in the operating environment of the gas turbine blade. There is preferably no non-ceramic coating on the lateral surface of the repaired blade tip.

In conventional practice, the recoating of the lateral surface of the tip of the turbine blade after the repair of the blade tip is exacting and involves several expensive and time-consuming steps. In the usual case where the coating includes a simple, non-ceramic aluminide environmental or bond coat, an aluminum-rich layer is deposited onto the base metal of the repaired area of the lateral surface of the turbine blade by a relatively slow process. The aluminum-rich layer is interdiffused into the base metal with an extended heat treatment. If the coating is a more complex aluminide such as a platinum aluminide, further processing is required to deposit and heat-treat diffuse a platinum layer prior to the deposition and heat-treat diffusion of the aluminum layer.

The present approach avoids the need to perform the several steps of the recoating process for the non-ceramic coating. The cost and processing time of the repair are thereby significantly reduced, making the repair a more attractive option in the decision whether to repair the damaged turbine blade or to install an expensive new-make turbine blade. The performance of the turbine blade repaired according to the present approach is acceptable, both for mechanical properties and for environmental resistance.

Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.