US20060219329A1

US20060219329A1 - Repair nickel-based superalloy and methods for refurbishment of gas turbine components

Info

Publication number: US20060219329A1
Application number: US11/093,350
Authority: US
Inventors: Yiping Hu; Richard Bye
Original assignee: Honeywell International Inc
Current assignee: Honeywell International Inc
Priority date: 2005-03-29
Filing date: 2005-03-29
Publication date: 2006-10-05

Abstract

A nickel-based superalloy includes, in terms of weight, about 0.06% to about 0.10% carbon, about 6.0% to about 6.4% aluminum, about 5.8% to about 6.3% tantalum, about 6.5% to about 7.0% chromium, about 8.8% to about 9.3% cobalt, about 0.6% to about 1.0% molybdenum, about 2.4% to about 2.8% rhenium, about 4.8% to about 5.3% tungsten, about 0.3% to about 0.80% hafnium, about 0.01% to about 0.03% zirconium, about 0.10% to about 0.18% silicon, and nickel. A method for repairing a surface of a turbine component includes the step of applying the nickel-based superalloy to a damaged area of the component surface. Post-deposition processes are also performed as necessary.

Description

TECHNICAL FIELD

The present invention relates to turbine engine components that function in high temperature and high pressure environments. More particularly, the present invention relates to methods for depositing a superalloy material onto turbine engine components to improve erosion resistance due to corrosion, oxidation, thermal fatigue, and other hazards.

BACKGROUND

Turbine engines are used as the primary power source for various kinds of aircrafts. The engines are also auxiliary power sources that drive air compressors, hydraulic pumps, and industrial gas turbine (IGT) power generation. Further, the power from turbine engines is used for stationary power supplies such as backup electrical generators for hospitals and the like.
Most turbine engines generally follow the same basic power generation procedure. Compressed air is mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turbine vanes in the engine. The vanes turn the high velocity gas flow partially sideways to impinge on the turbine blades mounted on a rotatable turbine disk. The force of the impinging gas causes the turbine disk to spin at high speed. Jet propulsion engines use the power created by the rotating turbine disk to draw more air into the engine and the high velocity combustion gas is passed out of the gas turbine aft end to create forward thrust. Other engines use this power to turn one or more propellers, electrical generators, or other devices.
Many turbine engine blades and vanes are fabricated from high temperature materials such as a nickel-based or cobalt-based superalloy. Although nickel-based and cobalt-based superalloys have good high temperature properties and many other advantages, they are susceptible to corrosion, oxidation, thermal fatigue and erosion damage in the high temperature environment of an operating turbine engine. These limitations are problematic as there is a constant drive to increase engine operating temperatures in order to increase fuel efficiency and to reduce emission. Replacing damaged turbine engine components made from nickel-based and cobalt-based superalloys is expensive, and significant research is undergone to find cost-effective ways to repair the components.
Hence, there is a need for methods and materials for repairing turbine engine components such as the turbine blades and vanes. There is a particular need for environment-resistant repair materials that will improve a turbine component's durability, and for efficient and cost effective methods of repairing the components using such materials.

BRIEF SUMMARY

The present invention provides a nickel-based superalloy. The superalloy comprises, in terms of weight percentage, about 0.06% to about 0.10% carbon, about 6.0% to about 6.4% aluminum, about 5.8% to about 6.3% tantalum, about 6.5% to about 7.0% chromium, about 8.8% to about 9.3% cobalt, about 0.6% to about 1.0% molybdenum, about 2.4% to about 2.8% rhenium, about 4.8% to about 5.3% tungsten, about 0.30% to about 0.80% hafnium, about 0.01% to about 0.03% zirconium, about 0.10% to about 0.18% silicon, 0.008% to 0.015% yttrium, and nickel.
The present invention also provides a method for repairing a surface of a turbine component. The method comprises the step of applying the nickel-based superalloy to a damaged area of the component surface. Other independent features and advantages of the preferred methods and nickel-based superalloys will become apparent from the following detailed description, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an exemplary turbine blade in accordance with an exemplary embodiment;
FIG. 2 is a flow diagram of a repair method in accordance with an exemplary embodiment;
FIG. 3 is a magnified view of a nickel-based superalloy powder according to one aspect of the present invention;
FIG. 4 is a plot of cyclic oxidation test results for a composition of the present invention and two nickel-based superalloys in common use; and
FIG. 5 is an optical micrographic view of a laser weld made on a conventional nickel-based superalloy substrate according to one aspect of the present invention.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

The following detailed description of the invention is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background of the invention or the following detailed description of the invention.
The present invention provides an improved method for repairing high pressure turbine (HPT) components such as turbine blades and vanes to improve resistance to erosion due to corrosion, oxidation, thermal fatigue, and other hazards. The alloy can be used to improve the ability of components such as turbine blades and vanes to resist erosion due to factors such as oxidation, corrosion, and thermal fatigue cracks, to name several examples.
The repair material is a nickel-based superalloy that is broadly defined as comprising carbon, aluminum, tantalum, chromium, cobalt, molybdenum, rhenium, tungsten, hafnium, boron, zirconium, silicon, and yttrium. One exemplary nickel-based superalloy includes each of these elements with a concentration (weight %) of carbon ranging between about 0.06% and about 0.10%, aluminum ranging between about 6.00% and about 6.40%, tantalum ranging between about 5.80% and about 6.30%, chromium ranging between about 6.50% and about 7.00%, cobalt ranging between about 8.80% and about 9.30%, molybdenum ranging between about 0.60% and about 1.00%, rhenium ranging between about 2.40% and about 2.80%, tungsten ranging between about 4.80% and about 5.3%, hafnium ranging between about 0.30% and about 0.80%, boron ranging between about 0.008% and about 0.014%, zirconium ranging between about 0.01% and about 0.03%, silicon ranging between about 0.10% and about 0.18%, and yttrium ranging between about 0.008% and about 0.015%.

In a preferred embodiment the combination of aluminum and tantalum is at least about 12.00% of the superalloy, and the combination of molybdenum, tungsten, and rhenium is at least about 8.00% of the superalloy. A more preferred embodiment includes carbon at about 0.08%, aluminum at about 6.20%, tantalum at about 6.00%, chromium at about 6.75%, cobalt at about 9.00%, molybdenum at about 0.80%, rhenium at about 2.60%, tungsten at about 5.00%, hafnium at about 0.660%, boron at about 0.01%, zirconium at about 0.02%, silicon at about 0.14%, and yttrium at about 0.01%. The acceptable concentrations of these elements, along with preferred embodiments, are presented in Table 1. In all of the various embodiments, the balance of the concentration is nickel.

TABLE 1


	[wt. %]	Preferred [wt. %]	More Preferred
Element	Range	Range	[wt. %]

C	0.06-0.10	0.06-0.10	0.08
Al	6.00-6.40	6.00-6.40	6.20
Ta	5.80-6.30	5.80-6.30	6.00
Al + Ta		≧12.00	12.20
Cr	6.50-7.00	6.50-7.00	6.75
Co	8.80-9.30	8.80-9.30	9.00
Mo	0.60-1.00	0.60-1.00	0.80
Re	2.40-2.80	2.40-2.80	2.60
W	4.80-5.25	4.80-5.25	5.00
Mo + W + Re		≧8.00	8.40
Hf	0.40-0.80	0.40-0.80	0.60
Bo	0.008-0.014	0.008-0.014	0.01
Zr	0.01-0.03	0.01-0.03	0.02
Si	0.10-0.18	0.10-0.18	0.14
Y	0.008-0.015	0.008-0.015	0.01
Ni	balance	balance	balance

As previously mentioned, the superalloy of the present invention can be used to restore both geometry and dimension to a variety of different turbine engine components. For example, the turbine blades in the hot section of a turbine engine are particularly susceptible to wear, oxidation and other degradation. One exemplary turbine blade that is repairable using the superalloy of the present invention is made from high performance Ni-based superalloys such as IN738, IN792, C101, MarM247, Rene80, Rene125, ReneN5, SC180, CMSX 4, and PWA1484.
Turning now to FIG. 1, a blade 150 that is exemplary of the types that are used in turbine engines is illustrated, although turbine blades commonly have different shapes, dimensions and sizes depending on gas turbine engine models and applications. The blade 150 includes several components that are particularly susceptible to erosion, wear, oxidation, corrosion, cracking, or other damage, and the process of the present invention can be tailored to deposit the nickel-based superalloy on different type of blades and their individual components. Among such blade components is an airfoil 152. The airfoil 152 includes a concave face and a convex face. In operation, hot gases impinge on the concave face and thereby provide the driving force for the turbine engine. The airfoil 152 includes a leading edge 162 and a trailing edge 164 that encounter air streaming around the airfoil 152. The blade 150 also includes a tip 160. In some applications the tip may include raised features commonly known as squealers. The turbine blade 150 is mounted on a non-illustrated turbine hub or rotor disk by way of a dovetail 154 that extends downwardly from the airfoil 152 and engages with a slot on the turbine hub. A platform 156 extends longitudinally outwardly from the area where the airfoil 152 is joined to the dovetail 154. A number of cooling channels desirably extend through the interior of the airfoil 152, ending in openings 158 in the surface of the airfoil 152.
As mentioned previously, the process of the present invention can be tailored to fit the blade's specific needs, which depend in part on the blade component where degradation has occurred. For example, the airfoil tip 160 is particularly subject to degradation due to oxidation, erosion, thermal fatigue and wear, and the repair process is used to apply the nickel-based superalloy of the present invention onto a machined airfoil tip 160. Other turbine airfoil components that are particularly subject to degradation include the knife edge seal, the leading and trailing edges 162, 164, and the platform 156.
It is also emphasized again that turbine blades are just one example of the type of turbine components that can be coated using the nickel-based superalloy of the present invention. Vanes, shrouds, and other turbine components can be repaired in the same manner according to the present invention.
A variety of different systems and implementations can be used to perform a repair method utilizing the inventive nickel-based superalloy. An exemplary repair method utilizes a welding system to restore a degraded airfoil component to a pre-designed geometry and dimension. Many different welding methods can be used to repair a metal component using the nickel-based superalloy of the present invention, including laser, plasma transfer arc (PTA), micro plasma, and tungsten inert gas (TIG) welding methods. Surface repair can also be performed using thermal spray process such as high velocity oxygen fuel (HVOF) thermal spraying, and low pressure plasma spraying (LPPS) methods.
Another process that can be utilized to perform a surface repair process is cold gas-dynamic spraying. In this process, the cold spray system mixes the repair particles with a suitable pressurized gas in a mixing chamber. The particles are then accelerated through the specially designed nozzle and directed toward a target surface on the turbine component. When the particles strike the target surface, the kinetic energy of the particles is converted into plastic deformation of the particle, causing the particle to form a strong bond with the target surface and produce a coating. In this method, the particles are applied and deposited in the solid state, i.e., at a temperature which is considerably lower than the melting point of the powder material. This system is but one example of the type of system that can be adapted to cold spray powder materials and to thereby repair a target surface.
Yet another process that can be utilized to perform a surface repair process is named a JetFix process, which is an activated diffusion brazing repair. An exemplary JetFix process includes cleaning the cracks with a hydrogen fluoride ion to remove oxides, applying the invented nickel-based superalloy and low-melt alloy mixture, vacuum furnace brazing, and final machining.
Turning now to FIG. 2, an exemplary method 200 is illustrated for repairing turbine blades, vanes, and other turbine components in a flow diagram. Although the following method is described with reference to repair of a turbine blade, it should be understood that the method is in no way limited to blades or any other particular components.
When one or more worn or damaged turbine blades are identified, they are typically detached from the turbine. The first step 202 of the method comprises preparing the surface on the turbine component. For example, the first step of preparing a turbine blade can include one or more processes including strip coating, pre-welding machining, degreasing and grit blasting the surface to be repaired in order to remove any oxidation and dirty materials. It will be appreciated that the present embodiment is not limited to these preparatory steps, and that additional, or different types and numbers of preparatory steps can be conducted.
Once the turbine blade has been prepared, the next step 204 comprises subjecting the blade to a repair process. The repair process can be a welding process to deposit the nickel-based superalloy material of the present invention onto the worn or damaged area, or other surface restoration processes such as thermal spraying, and cold gas-dynamic spraying repair. The nickel-based superalloy of the present invention can be deposited onto the worn or damaged nickel-based or cobalt-based turbine blade area using any of these repair processes because the superalloy of the present invention typically has mechanical properties that allow the welded metals to be compatible with the substrate. In an exemplary embodiment, the nickel-based superalloy has mechanical properties that substantially match or improve upon those of the turbine blade superalloy.
An exemplary repair process includes laser welding with the nickel-based superalloy of the present invention. Laser welding is particularly effective when the nickel-based superalloy filler material is provided as substantially spherical powder particles since spherical powders can improve powder flow property and help maintain a stable powder feed rate during the welding process. FIG. 3 is a magnified view of the nickel-based superalloy powder having substantially spherical particle morphology.
Laser welding can be carried out using a CO₂laser, a YAG laser, a diode laser, or a fiber laser. One exemplary welding process includes laser powder fusion welding, in which the nickel-based superalloy of the present invention is laser deposited onto a degraded area to restore both geometry and dimension with metallurgically sound buildup. Both automatic and manual laser welding systems are widely used to perform laser powder fusion welding processes. An exemplary manual welding repair is described in detail in U.S. Pat. No. 6,593,540 entitled “Hand Held Powder-Fed Laser Fusion Welding Torch” and incorporated herein by reference.
Returning to the flow diagram of FIG. 2, after the repair step 204 is completed at least one post-repair step 206 is performed depending on the type of repair that was performed. For example, if a spraying process was performed to repair the blade 150, then one exemplary post-repair step is a hot isostatic pressing (HIP) process performed for about four hours at 2200° F. and with an applied pressure of 15 Ksi.
The post-deposition step 206 can further include additional processes that improve the turbine blade's mechanical properties, and metallurgical integrity. Such processes include final machining the repaired components to a predetermined design dimension. Other processes include re-coating with a suitable material such as environment-resistant diffusion aluminide and/or MCrAlY overlay coatings, coating diffusion, and aging heat treatments.
After the post-deposition step 206 is completed, at least one inspection process can be performed as step 208 to determine whether any surface defects exist, such as cracks or other openings. An inspection process can be conducted using any well-known non-destructive inspection techniques including, but not limited to, a fluorescent penetration inspection, and a radiographic inspection. If an inspection process indicates that surface defects do exist, the turbine blade 150 is subjected to an additional repair process. This additional process may be either another laser welding process or a diffusion brazing process.
The present invention thus provides a novel nickel-based superalloy and an improved method for repairing turbine engine components. The repair method restores the components and improves their durability, thereby optimizing the operating efficiency of a turbine engine, and prolonging the operational life of turbine blades and other engine components.
The following examples are presented in order to provide a more complete understanding of the invention. The specific techniques, conditions, materials and reported dada set forth to illustrate the invention are exemplary and should not be construed as limiting the scope of the invention.

EXAMPLE 1

An alloy, having a nominal composition of 6.2 wt. % Al, 9 wt. % Co, 6.75 wt. % Cr, 0.45 wt. % Hf, 0.8 wt. % Mo, 2.6 wt. % Re, 0.15. wt % Si, 6.0 wt. % Ta, 5 wt. % W, 0.02 wt. % Zr, 0.01 wt. % Y, 0.01 wt. B, and 0.08 wt. % C, with the balance being Ni plus incidental impurities, was vacuum induction melted and cast. The cast ingot was hot isostatically pressed at a pressure of 15 ksi and a temperature of 2200° F. for four hours.
Pieces measuring approximately 0.5 in.×0.5 in.×0.125 in. were machined from the ingot. The surfaces of the samples were sanded with 320 grit paper and lightly grit blasted. The pieces were exposed to repeated cycles in which they were alternatively placed in a furnace chamber maintained at 2100° F. for twenty-seven minutes, removed and cooled in moving room temperature air for three minutes, and then reinserted into the 2100° F. furnace for the next cycle. The samples were periodically removed and weighed in order to determine the weight change as a function of the number of cycles.
Similarly sized pieces of two superalloys commonly used for gas turbine engine components were prepared in the same manner and tested alongside the samples of the invention for comparison purposes. The two superalloys are CMSX4 (nominal composition: 5.6 wt. % Al, 9.0 wt. % Co, 6.3 wt. % Cr, 0.1 wt. % Hf, 0.6 wt. % Mo, 3.0 wt. % Re, 6.5 wt. % Ta, 6.0 wt. % W, 1.0 wt. % Ti, with the balance being Ni plus incidental impurities), and MarM 247 (nominal composition: 5.5 wt. % Al, 10.0 wt. % Co, 8.2 wt. % Cr, 1.5 wt. % Hf, 0.6 wt. % Mo, 3.0 wt. % Ta, 10.0 wt. % W, 1.0 wt. % Ti, 0.05 wt. % Zr, 0.015 wt. % B, and 0.16 wt. % C, with the balance being Ni plus incidental impurities).
The results of the cyclic oxidation testing is shown in FIG. 4, in which weight change divided by the original sample surface area is plotted against the number of cycles. The plotted lines represent the average values from multiple samples of each composition. The composition of the invention demonstrates significantly better cyclic oxidation behavior than either of the two comparison compositions.

EXAMPLE 2

Metal powders were produced via the plasma rotating electrode process from the cast ingot referred to in Example 1. Using an automatic laser welding system, the powders were fed to the molten pool coaxially with a laser beam to form a clad on the substrate. Three overlapping weld beads of a composition of the present invention were made on an advanced nickel-based superalloy substrate that was approximately 0.060 in. wide to simulate a turbine blade tip repair. The welded sample was subsequently heat treated, cross sectioned and examined metallographically. FIG. 5 is a photomicrograph of the weld cross section, demonstrating a metallurgically sound weld resulting from the optimized welding parameters.
While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt to a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

1. A nickel-based superalloy, comprising in terms of weight:

about 0.06% to about 0.10% carbon;

about 6.0% to about 6.4% aluminum;

about 5.8% to about 6.3% tantalum;

about 6.5% to about 7.0% chromium;

about 8.8% to about 9.3% cobalt;

about 0.6% to about 1.0% molybdenum;

about 2.4% to about 2.8% rhenium;

about 4.8% to about 5.3% tungsten;

about 0.3% to about 0.8% hafnium;

about 0.01% to about 0.03% zirconium;

about 0.10% to about 0.18% silicon; and

nickel.

2. The nickel-based superalloy of claim 1, further comprising in terms of weight:

about 0.008% to about 0.014% boron; and

about 0.008% to about 0.015% yttrium.

3. The nickel-based superalloy of claim 1, wherein in terms of weight at least 12% is aluminum and tantalum combined.

4. The nickel-based superalloy of claim 1, wherein in terms of weight at least 8.0% is molybdenum, tungsten, and rhenium combined.

5. The nickel-based superalloy of claim 1, wherein the superalloy is prepared in the form of a powder.

6. The nickel-based superalloy of claim 1, wherein the superalloy is a prepared in the form of a weld wire.

7. The nickel-based superalloy of claim 5, wherein the powder comprises substantially spherical particles.

8. A method for repairing a surface of a turbine component, comprising the step of:

applying a nickel-based superalloy to a damaged area of the component surface, the nickel-based superalloy comprising in terms of weight:

about 0.06% to about 0.10% carbon;

about 6.0% to about 6.4% aluminum;

about 5.8% to about 6.3% tantalum;

about 6.5% to about 7.0% chromium;

about 8.8% to about 9.3% cobalt;

about 0.6% to about 1.0% molybdenum;

about 2.4% to about 2.8% rhenium;

about 4.8% to about 5.3% tungsten;

about 0.3% to about 0.8% hafnium;

about 0.01% to about 0.03% zirconium;

about 0.10% to about 0.18% silicon; and

nickel.

9. The method of claim 8, wherein the nickel-based superalloy further comprises in terms of weight:

about 0.008% to about 0.014% boron; and

about 0.0080% to about 0.015% yttrium.

10. The method of claim 8, wherein in terms of weight at least 12% of the nickel-based superalloy is aluminum and tantalum combined.

11. The method of claim 8, wherein in terms of weight at least 8.0% of the nickel-based superalloy is molybdenum, tungsten, and rhenium combined.

12. The method of claim 8, wherein the nickel-based superalloy is prepared in the form of a powder.

13. The method of claim 8, wherein the nickel-based superalloy is prepared in the form of a weld wire.

14. The method of claim 12, wherein the powder comprises substantially spherical particles.

15. The method of claim 8, wherein the nickel-based superalloy is applied to the damaged area while performing a welding process on the damaged area.

16. The method of claim 8, wherein the nickel-based superalloy is applied to the damaged area using a cold gas dynamic spraying process.

17. The method of claim 8, wherein the nickel-based superalloy is applied to the damaged area using a thermal spraying process.

18. The method of claim 17, wherein the thermal spraying process is selected from the group of processes consisting of high velocity oxygen fuel thermal spraying, and low pressure plasma spraying.

19. The method of claim 8, further comprising the step of:

subjecting the component to a hot isostatic pressing process after applying the nickel-based superalloy to the damaged area.

20. The method of claim 8, further comprising the step of:

heat treating the turbine component after applying the nickel-based superalloy to the damaged area.