Patent Application
20160010472
The present invention relates to a protective coating for a gas turbine blade or other component wherein the coating includes an aluminum-bearing coating applied at a relatively high temperature region of the component and a chromium-bearing coating applied at another relatively lower temperature region of the component depending on coating functionality needed.
Current gas turbine designs are requiring that a variety of coatings be applied to different areas of the turbine part for different functional reasons. Examples of coating functionality include wear, oxidation, thermal barrier, and hot corrosion. Turbine designers choose an appropriate coating for a particular functionality in the gas turbine environment.
Hot corrosion is a form of accelerated oxidation when a liquid salt is present on the surface of a Ni and Co based superalloy component. The salt is usually sodium sulfate with other naturally occurring constituents, such as K, Ca, and/or Mg, present. It is well known that as the Cr content of an alloy increases, its resistance to hot corrosion attack increases. Current methods to increase surface Cr content are pack and vapor phase chromizing, which comprise one-step deposition and reaction with the Ni substrate alloy, forming a Cr-enriched alloy zone. The chromizing process is facilitated by halide (Cl or F) activators that form Cr-halide gases at relatively high temperatures, such as greater than 1900 degrees F.
Since pack and vapor phase chromizing require high temperature application above 1900 F and are difficult to apply to localized part areas of interest, these processes must be applied early in the part routing to the entire the part. Masking has not been effective in these processes as a means for controlling the localized deposition of the chromium on certain areas of interest and, as a result, has not been applied in these high temperature processes.
The present invention provides in an embodiment a method of forming a protective coating on a gas turbine component wherein the duplex coating includes an aluminum-bearing coating applied at one region of the gas turbine component where relatively higher temperatures are encountered in service and a chromium-bearing coating applied at another region of the turbine blade or other component where relatively lower temperatures and hot corrosion are encountered in service, thereby providing coating functionality for the different temperatures and oxidation/corrosion environments to be encountered by the gas turbine component.
In an illustrative embodiment of the present invention, the method involves forming a duplex coating on a superalloy substrate by first applying an aluminum-bearing coating on the first relatively higher temperature region of the substrate, secondly applying a metallic coating comprising chromium on an adjacent relatively lower temperature region of the substrate followed by diffusing chromium into the substrate to form a chromium-enriched diffused coating thereon at the adjacent relatively lower temperature region. The aluminum-bearing coating is applied in a first step by high temperature vapor deposition, while the chromium-bearing coating is applied in a subsequent second step at a relatively lower temperature, such as less than 500° F. The method typically involves applying masking on the relatively lower temperature region before the aluminum-bearing coating is applied on the relatively higher temperature region and subsequently applying masking on the relatively higher temperature region before the metallic coating of chromium is applied on the relatively lower temperature region.
In the event the substrate is a gas turbine component, the method is practiced by first applying a mask on a root region of the component, then applying an aluminum-bearing coating, such as a diffusion aluminide coating, on an airfoil region, de-masking the root portion, and then masking the already-coated airfoil region. Then, the method involves depositing a metallic coating comprising chromium on at least a portion of a relatively lower temperature root region that will be subject to hot corrosion, de-masking the airfoil region followed by diffusing the chromium into the alloy at the coated portion of the root region to form a chromium-enriched diffused surface coating on the portion of the root region. The aluminum-bearing coating optionally can be applied to cover the airfoil region and also an intermediate platform region and root shank region. An attachment portion, such as a fir tree portion, of the root region may be left uncoated to enhance fatigue life of the root region where it is connected to a turbine disk.
In a particular embodiment of the present invention, a relatively low temperature deposition process embodying a liquid deposition medium, such as electroplating bath, electrophoretic bath, liquid slurry, and others, is used to form a metallic coating comprising a majority of chromium on at least a portion of the relatively lower temperature region of a precursor component. The chromium coating is applied as a very thin layer having a thickness of 0.00005 to 0.005 inch. Diffusion of the as-deposited chromium into the substrate typically is effected by elevated temperature heat treatment after the masking is removed from the previously-applied aluminum-bearing coating on the airfoil region.
The present invention envisions a nickel or cobalt based alloy turbine component precursor having an aluminum-bearing coating applied on an airfoil region and metallic coating comprising substantially pure chromium or a chromium alloy applied on at least a portion of the root region of the component. The chromium coating then is diffused into the alloy to form a diffused chromium-enriched coating on the portion of the root region of the gas turbine blade. The diffused chromium-enriched coating has an outermost region that comprises at least about 20%, preferably about 25%, and more preferably about 30% to about 60% by weight Cr.
Advantages, features, and embodiments of the present invention will become apparent from the following description.
In one embodiment of the present invention, a method is provided for forming a protective coating on a gas turbine component wherein the coating includes an aluminum-bearing coating applied at one region of the turbine blade or other component where relatively higher temperatures are encountered in service and a chromium-bearing coating formed at another adjacent region of the component where relatively lower temperatures and hot corrosion are encountered in service. Such a duplex coating provides coating functionality for the different temperatures and oxidation/corrosion environments to be encountered in service.
The present invention is especially useful for protecting different regions of a gas turbine blade component from oxidation and hot corrosion in service in a gas turbine engine, although the invention is not limited to gas turbine components since it can be practiced to protect other components against oxidation and hot corrosion. The present invention can be practiced to protect nickel based superalloy gas turbine components, nickel-cobalt based superalloy gas turbine components, or cobalt based superalloy gas turbine components from hot corrosion, although the invention is not limited to these alloys. For purposes of illustration and not limitation, the present invention will be described below with respect to protection of different regions of a gas turbine engine blade made of CMSX-4 nickel based superalloy against oxidation and hot corrosion in service in a gas turbine engine.
In particular,
A first hotter region of the turbine blade 10 is subjected to relatively higher temperatures and oxidation degradation in service in the gas turbine engine and comprises the airfoil region 12 and surface 20a of a platform region 20 that faces toward the airfoil region such that the airfoil region 12 and platform surface 20a operate in or near the hot gas path of the turbine section of the gas turbine engine. The airfoil region 12 and platform surface 20a are the hottest regions of the turbine blade and usually operate above 1900 degrees F. for purposes of illustration and not limitation.
As a result of the relatively high operating temperatures encountered, the airfoil region 12 and the platform surface 20a preferably are provided with a so-called alumina-former coating thereon that produces an adherent protective scale of alumina in service in the gas turbine engine.
A second relatively cooler region of the turbine blade 10 is subjected to relatively lower temperatures and hot corrosion by salts, such as sodium sulfates and other constituents such as K, Ca, and/or Mg, in service in the gas turbine engine. The second region comprises the under (lower) surface 20b of a platform region 20 that faces away from the airfoil region 12 and the root region 14. The second region thus involves a cooler region that on older turbine blades may operate uncoated. However, as the combustor efficiency has improved, the operating temperature of the second region is generally increasing and spread more uniformly over the second region. Hence the first region comprised of the airfoil 12 and platform surface surface 20a is also becoming hotter. When salts of sodium sulfate are deposited on a surface that operates between 1200 F and 1850 F, hot corrosion attack can occur. Combining the high stress state of the blade root 14, with hot corrosion conditions, rapid attack and fracture of the turbine blade in the root region can occur. For turbine blades with uncoated root regions heretofore used in the lower temperature operating conditions, hot corrosion resistance can be increased by increasing the chromium content of the turbine blade alloy.
The present invention provides a multiplex coating and a method for applying the coating to protect the different hotter and cooler regions of the turbine blade exposed to more aggressive temperature/hot corrosion conditions associated recent engine designs. The present invention provides a nickel or cobalt based alloy turbine blade 10 having an aluminum-bearing coating AL applied on an airfoil region of the blade and a metallic coating comprising chromium applied on at least a portion of the root region of the blade and diffused into the alloy to form the diffused chromium-enriched coating CR on the portion of the root region.
The aluminum-bearing coating is applied in a first one step procedure by high temperature vapor deposition, such as by chemical vapor deposition at or above 1900 degrees F. pursuant to U.S. Pat. Nos. 5,264,245; 4,132,816; and 3,486,927, by conventional above-the-pack processes, or other vapor deposition processes.
The chromium-bearing coating is applied after the aluminum-bearing coating using a two step procedure that involves depositing a metallic coating comprising chromium on the substrate at a relatively low temperature below 212 degrees F. when a liquid electrolytic deposition bath or liquid carrier medium is employed followed by a high temperature heat treatment to diffuse chromium into the substrate. Exemplary low temperature processes for depositing the metallic chromium coating include, but are not limited to, electroplating or electrophoetric deposition using a liquid bath, and slurry coating with chromium-bearing particles (e.g. Cr or Cr alloy particles) in a liquid carrier followed by drying, all of which can be conducted below 212 degrees F. using liquid baths or liquid slurries. Certain other relatively low temperature deposition processes can be employed to deposit the metallic coating comprising chromium including, but not limited to, electro-spark discharge conducted typically at less than 500° F., cladding conducted typically at less than 100° F., plasma spray conducted at less than 500° F., and entrapment plating wherein Cr particles are entrapped in a Ni electroplated layer.
When the substrate comprises a gas turbine component having airfoil, platform and root regions, a method embodiment is practiced by first applying a mask on a root region of the component, then applying an aluminum-bearing coating, such as a diffusion aluminide coating, on an airfoil region, de-masking the root portion, and then masking on the already-coated airfoil region. Then, this method embodiment deposits a metallic coating comprising chromium on at least a portion of a relatively lower temperature root region that will be subject to hot corrosion, de-masks the airfoil region, followed by diffusing the chromium into the alloy at the coated portion of the root region to form a chromium-enriched surface coating on the portion of the root region. The aluminum-bearing coating optionally can be applied to cover the airfoil region and also an intermediate platform region and root shank region. An attachment portion, such as a fir tree portion, of the root region may be left uncoated to enhance fatigue life of the root region where it is connected to a turbine disk.
The chromium-enriched diffused coating applied on at least a portion of the root region of a nickel base superalloy substrate typically comprises in the diffused condition a Cr-enriched outermost diffusion zone comprising chromium, nickel, and other substrate alloy elements in solid solution wherein Cr is present as a majority of the zone,
Practice of embodiments of the invention allow control of the Cr content and Cr depth profile into the substrate to tailor hot corrosion protection as needed for a particular service application. Typically, more Cr at the outermost coated substrate surface will be more protective than less. More Cr can be provided by varying the thickness of the Cr metallic coating and the diffusion heat treatment conditions.
This Comparison Example is offered to help illustrate the problems and difficulties of forming such a duplex coating on a gas turbine blade by processing other than that pursuant to the present invention.
For example, available high temperature (above 1900 F) pack or vapor phase chromizing processes and high temperature (above 1700 degrees F.) pack, vapor phase or CVD aluminizing processes can be used to produce a turbine blade with two environmental protection coatings (duplex coating). However problems in processing and retaining high surface Cr have been observed. The masking used for aluminizing can remove Cr from the chromized shank during the high temperature aluminizing process. Namely, in order to coat a turbine blade with the duplex coating, the turbine blade must be entirely chromized by a high temperature pack or vapor phase process and then the resulting Cr-rich layer must be removed from the gas path surfaces 12, 20a prior to aluminizing or overcoating the gas path surfaces. To prevent aluminizing of the root region 14, the root region is masked by placing it in masking powder (e.g. alumina powder, NiO powder, etc.) residing in a containment box. However, this procedure has resulted in unwanted reductions in Cr content of the previously applied Cr-enriched layer on the root region and a reduction in its hot corrosion resistance as will now be demonstrated.
A cast turbine blade having airfoil, platform, and root features of
The chromizing process was conducted using the following pack parameters: pure chromium powder with aluminum oxide and NH4Cl activator for 5 hours at 1950 degrees F.
The Pt electroplating was conducted using the following parameters set forth in U.S. Pat. No. 5,788,823, which is incorporated herein by reference to this end, to deposit 0.3 mils of Pt on the substrate. The aluminizing process was conducted using the following parameters: 1975 degrees F. for 1440 minutes in H2/AlCl3 atmosphere pursuant to U.S. Pat. No. 5,264,245, which is incorporated herein by reference to this end.
The enriched Cr content of the as-chromized coating on the second region as shown in
Pursuant to method embodiments of the present invention, the duplex coating is applied using a sequence processing steps that overcomes the above-discussed problems and difficulties demonstrated in the Comparison Example.
Pursuant to an illustrative embodiment of the present invention, the following processing steps are employed:
1. If a platinum-modified diffusion aluminide coating is to be formed on the gas path surfaces 12, 20a, then these surfaces are optionally electroplated with a layer of Pt pursuant to U.S. Pat. No. 5,788,832 which is already incorporated herein by reference. If a simple diffusion aluminde coating is to be formed, then this step is omitted.
2. Masking the second region of the turbine blade (i.e. root region 14 and platform surface 20b) with the M1 maskant powder mentioned above in a containment box. That is, the root region 14 and platform surface 20b are embedded in the maskant powder in the containment box.
3. Aluminize the first hotter region (i.e. airfoil 12 and platform surface 20a) to form a diffusion aluminide coating, such as a Pt-modified diffusion aluminide coating if step 1 is practiced, with the masking covering the second region.
4. Masking the diffusion aluminide coating on the first region.
5. Cr electroplating the second cooler region with the masking of step 2 covering the diffusion aluminide coating formed in step 3. The Cr electroplating is conducted at low temperature such as less than 212 degrees F. using a liquid (e.g. aqueous) electroplating bath. The Cr electroplate can be locally deposited by virtue of the masking on the first region being effective under the low temperature plating bath conditions.
6. Diffusing the Cr plating into the CMSX-4 substrate alloy to form the Cr-enriched hot corrosion resistant coating wherein diffusing of the Cr plating improves bonding with the superalloy substrate and makes the resulting Cr-rich layer more ductile.
The chromium electroplating process is conducted using plating conditions to deposit a hexavalent hard, dense chromium electroplate comprising substantially pure Cr that meets AMS (Aerospace Material Specification) 2438B for hard, dense chromium coatings for aerospace material applications on steel materials. AMS 2338B is incorporated herein by reference to this end.
In this example, the hard, dense substantially pure chromium electroplate was applied commercially by a commercial electroplater Armoloy of Illinois, 118 Simonds Ave., DeKalb, Ill., using proprietary plating conditions. The deposited Cr electroplating was applied to a thickness of 8.7 micrometers or 3.5 micrometers. The electroplated layer was substantially pure Cr; e.g. 99.9% by weight pure Cr and balance plating impurities. The invention envisions electroplating Cr alloys, rather than pure Cr, and also plating alternating layers of Cr and Ni.
The chromium electroplating can be conducted using any suitable parameters. For purposes of illustration and not limitation, the following plating conditions can be used:
1. Vapor hone surfaces with an alumina slurry to clean surfaces to be plated.
2. Activate the surfaces to be plated by immersion in plating bath containing 250-400 g/L chromic acid and 2.5-4 g/L of sulfate catalyst (sulfuric acid) at 52-63° C. and applying a current (30-54 A/dm2 at 3 to 12 volts) such that the parts are anodes (which is opposite of Cr plate deposition) for 30 seconds to 2 minutes.
3. Cr plate surfaces to be plated by immersion in plating bath and applying current (such that the parts are cathodes) for 4 minutes to 30 minutes or as long as needed to meet the thickness requirement for the Cr plating.
4. Rinse in 120° F. de-ionized water to remove majority of plating bath.
5. Rinse in hot de-ionized water to remove remaining plating bath and dry.
The CVD aluminizing process is conducted using the following parameters: 1975 degrees F. for 1440 minutes in H2/AlCl3 atmosphere pursuant to U.S. Pat. No. 5,264,245, which is incorporated herein by reference to this end. Other aluminizing processes which can be used include, but are not limited to, pack, vapor phase, sputtering, physical vapor deposition and slurry followed by diffusion heat treatment, electrophoretic followed by diffusion heat treatment, and others.
For this example, the diffusion heat treatment of Cr was conducted at 1975 degrees F. for 4 hours in an Ar partial pressure atmosphere or at 2050 degrees F. for 2 hours in an Ar partial pressure atmosphere to prevent oxidation.
Comparing the later two samples in
Pursuant to another illustrative embodiment of the present invention, the following processing steps are employed:
1. If a platinum-modified diffusion aluminide coating is to be formed on the gas path surfaces 12, 20a, then these surfaces are optionally electroplated with a layer of Pt pursuant to U.S. Pat. No. 5,788,832 which is already incorporated herein by reference. If a simple diffusion aluminide coating is to be formed, then this step is omitted.
2. Aluminize the first hotter region and the second region to form a diffusion aluminide coating, such as a Pt-modified diffusion aluminide coating. No masking covering the second region.
3. Removing the diffusion aluminide coating selectively from the second region by grit blasting, machining or other technique to expose the substrate alloy, while leaving the diffusion aluminide coating on the first region.
4. Masking the diffusion aluminide coating on the first region as described in Example 1.
5. Cr electroplating the exposed second cooler region with the masking of step 4 covering the diffusion aluminide coating formed in step 2. The Cr electroplating is conducted at low temperature such as less than 212 degrees F. using a liquid electroplating bath. The Cr electroplate can be locally deposited by virtue of the masking on the first region being effective under the low room temperature plating bath conditions.
5. Diffusing the Cr plating into the CMSX-4 substrate alloy to form the Cr-enriched hot corrosion resistant coating wherein diffusing of the Cr plating improves bonding with the superalloy substrate and makes the resulting Cr-rich layer more ductile.
The chromium electroplating process is conducted by the commercial electroplater of Example 1. The CVD aluminizing process is conducted using the parameters of Example 1. The diffusion heat treatment of Cr is conducted using the parameters of Example 1.
The invention allows for many combinations of Al, Al/Cr, Cr, and bare, uncoated areas on a turbine blade to provide desired coating functionality as needed to suit different service conditions in the gas turbine engine.
Although the invention has been described with respect to certain detailed embodiments thereof, it will be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention.
This application claims benefit and priority of U.S. provisional application Ser. No. 61/633,935 filed Feb. 21, 2012, the entire disclosure of which is incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| 61633935 | Feb 2012 | US |
