Patent Application
20090274562
The present invention relates to the manufacture and restoration of aerospace components, such as components of gas turbine engines. In particular, the present invention relates to methods for forming coatings on turbine-stage nozzle segments.
Gas turbine engines operate by burning a combustible fuel-air mixture, and converting the energy of combustion into a propulsive force. A gas turbine engine typically includes an inlet, a compressor, a combustor, a turbine, and an exhaust duct, where the compressor draws in ambient air and increases its temperature and pressure. Fuel is added to the compressed air in the combustor, where it is burned to raise the gas temperature, thereby imparting energy to the gas stream. The resulting combustion gases are directed axially rearward from the combustor through an annular duct, where the gases interact with multiple turbine stages disposed within the annular duct.
Each turbine stage includes a stationary turbine nozzle derived of multiple stator vanes, and a downstream row of rotatable blades. The stator vanes direct the combustion gases axially rearward in a downstream direction, and the rotatable blades direct the energy of the combustion gases to an axial drive shaft that is interconnected with the compressor. Stator vanes typically have airfoil geometries designated by concave pressure sides and convex suction sides that extend axially between corresponding leading and trailing edges of the airfoils. Each airfoil is also typically disposed circumferentially between an outer arcuate shroud and an inner arcuate platform, thereby forming a nozzle segment. Multiple nozzle segments are interconnected to form the annular ring of the stationary turbine nozzle. Each nozzle segment may be cast to include one or more stator vanes disposed between the same outer arcuate shroud and inner arcuate platform. For example, a nozzle segment containing a single stator vane is typically referred to as a nozzle singlet, a nozzle segment containing two stator vanes is typically referred to as a nozzle doublet, nozzle segment containing three stator vanes is typically referred to as a nozzle triplet, and so on.
The components of the turbine stages (e.g., vanes and blades) are required to be able to withstand the thermal and oxidation conditions of the high temperature combustion gas during the course of operation. To protect turbine engine components from the extreme conditions, such components are typically coated with metallic bond coats that provide oxidation and/or corrosion resistance, and with ceramic thermal barrier coatings to provide thermal protection. However, many coating processes for forming bond coats and thermal barrier coatings require line-of-sight depositions. This poses problems when coating nozzle segments having multiple stator vanes because the inboard surfaces of the stator vanes are partially shadowed from the line-of-sight depositions. The resulting coatings typically have high variations in coating thicknesses due to the partial shadowing. Thus, there is an ongoing need for methods for coating nozzle segments having multiple stator vanes (e.g., nozzle doublets and triplets) to provide substantially uniform coatings with line-of-sight coating techniques.
The present invention relates to a coated nozzle segment having a plurality of stator vanes disposed between an outer shroud and an inner platform, and a method for coating the nozzle segment. The method includes separating the outer shroud and the inner platform along pathways that bisect first and second vanes of the plurality of stator vanes, coating the first and the second vanes after separating the outer shroud and the inner platform, and rejoining each of the separated outer shroud and the separated inner platform after coating the first and the second vanes.
After the outer shroud and the inner platform are each separated, the stator vanes of the sub-segments are coated using a variety of coating techniques to form one or more coatings on the exposed surfaces of the stator vanes (step 18). Suitable coatings for the stator vanes include metallic protective coats (e.g., bond coats), thermal barrier coatings, and combinations thereof. As discussed below, the separation of the outer shroud and the inner platform allows the stator vanes to be placed apart from each other during the coating process, thereby allowing each stator vane to be coated with a line-of-sight coating technique. In one embodiment, this allows the resulting coatings to have substantially uniform thicknesses. As used herein with reference to a coating on a surface, the term “substantially uniform thickness” refers to coating thicknesses along the surface (e.g., along a surface of a stator vane) that remain within 10% of an average coating thickness for the given coating on the surface, disregarding thickness deviations due to topographical variations in the surface (e.g., cooling holes in the surface). Alternatively, in embodiments in which an non-uniform coating thickness is desirable (e.g., controlled changes in thickness), the separation of the outer shroud and the inner platform allow each stator vane to be coated with the desired changes in coating thicknesses.
After the coating process is complete, the outer shroud and the inner platform are each rejoined using a suitable joining process. The joining process forms bond lines between the sub-segments that are desirably capable of withstanding the extreme temperatures and pressures of the turbine stages of a gas turbine engine. The joining process also desirably preserves the integrities of the previously applied coatings (steps 20 and 22). The resulting nozzle segment, having the coated stator vanes, may then be reassembled with additional nozzle segments to form an annular ring of a stationary turbine nozzle. As discussed below, the separation and rejoining of the outer shroud and the inner platform allows the coatings formed on the stator vanes to have substantially uniformly thicknesses, thereby preserving the oxidation, corrosion, and/or thermal resistances of the stator vanes during the course of operation in a gas turbine engine.
Similarly, vane 28 includes interior region 44, leading edge 46, pressure sidewall 48, and suction sidewall 50, where interior region 44 is a hollow interior portion of vane 28 that directs the flow of cooling air during operation. Stator vane 28 also includes a trailing edge (not shown in
Shroud 30 is an outer arcuate band secured to vanes 26 and 28, thereby allowing cooling air to enter interior regions 34 and 44 during operation. Shroud 30 includes leading edge 52, trailing edge 54, pressure side edge 56, and suction side edge 58, where leading edge 52 and trailing edge 54 are the upstream and downstream edges of shroud 30, respectively. Pressure side edge 56 and suction side edge 58 are the lateral edges of shroud 30, and are the edges that are secured to outer shrouds of adjacent nozzle segments (not shown) with leaf seal engagements to form a stationary turbine nozzle.
Correspondingly, platform 32 is an inner arcuate band secured to vanes 26 and 28, opposite of shroud 30. Platform 32 includes leading edge 60, trailing edge 62, pressure side edge 64, and suction side edge 66, where leading edge 60 and trailing edge 62 are the upstream and downstream edges of platform 32, respectively. Pressure side edge 64 and suction side edge 64 are the lateral edges of platform 32, and are the edges that are secured to inner platforms of adjacent nozzle segments (not shown) with leaf seal engagements to form the stationary turbine nozzle.
As shown, pressure sidewall 40 of vane 26 and suction sidewall 50 of vane 28 are outboard surfaces that are directly accessible with a line-of-sight coating technique. As a result, coatings may be readily deposited on pressure sidewall 40 and suction sidewall 50 with substantially uniform thicknesses. In contrast, however, suction sidewall 42 of vane 26 and pressure sidewall 48 of vane 28 are inboard surfaces, which partially shadow each other. The partial shadowing prevents line-of-sight coating techniques from evenly depositing coatings on suction sidewall 42 and pressure sidewall 48, thereby reducing coating thickness uniformity. The reduction in coating thickness uniformity correspondingly reduces the effectiveness of the formed coatings in providing corrosion, oxidation, and/or thermal resistance during the course of operation in a gas turbine engine.
As discussed above, method 10 is suitable for forming coatings on vanes 26 and 28, where the formed coatings have substantially uniform thicknesses on the outboard surfaces and the inboard surfaces. Prior to performing the coating process, pursuant to step 12 of method 10, pathways 68 and 70 are identified along shroud 30 and platform 32, respectively. Pathway 68 desirably extends from leading edge 52 to trailing edge 54 of shroud 30, between vanes 26 and 28. Similarly, pathway 70 desirably extends from leading edge 60 to trailing edge 62 of platform 32, also between vanes 26 and 28. As such, pathways 68 and 70 bisect vanes 26 and 28 along shroud 30 and shroud 32, respectively. While shown as linear pathways, pathways 68 and 70 may alternatively be non-linear pathways (e.g., curved lines and angled-segmented lines).
In one embodiment, pathways 68 and 70 are identified at locations along shroud 30 and platform 32 that are substantially even between vanes 26 and 28. This reduces the risk of damaging vanes 26 and 28 during the separation and rejoining steps of method 10. In an additional embodiment in which nozzle segment 24 includes existing bonds lines between vanes 26 and 28 (e.g., obtained during a previous manufacturing or restoration joining process), pathways 68 and 70 desirably follow the existing bond lines. This allows shroud 30 and platform 32 to be separated along the existing bond lines, which preserves the alloy microstructures of shroud 30 and platform 32.
After pathways 68 and 70 are identified, shroud 30 and platform 32 are separated along pathways 68 and 70, pursuant to steps 14 and 16 of method 10. Shroud 30 and platform 32 may be separated using a variety of techniques that are suitable for cutting the alloys of shroud 30 and platform 32 without damaging vanes 26 and 28. In one embodiment, shroud 30 and platform 32 are separated (simultaneously or sequentially) using wire electrical discharge machining (EDM). In this embodiment, nozzle segment 22 is placed in an aqueous bath, and a conductive wire is aligned with pathways 68 and 60 at leading edges 52 and 60, or at trailing edges 54 and 62. Electrical discharges are then sent through the conductive wire, thereby vaporizing successive portions of shroud 30 and platform 32 in the vicinity of the conductive wire. The wire is then passed along pathways 68 and 70 to separate nozzle segment 22 into separate sub-segments, thereby bisecting vanes 26 and 28.
After the separation process, sub-segments 24a and 24b are placed apart from each other, and vanes 26 and 28 are coated to form one or more protective coatings, pursuant to step 18 of method 10. Prior to the coating process, one or more surfaces of shroud portions 30a and 30b and platform portions 32a and 32b may be masked to prevent the formation of coatings on the masked surfaces. Vanes 26 and 28 may then be coated using a variety of coating techniques, including line-of-sight coating techniques. For metallic bond and protective coatings, suitable coating techniques include electron-beam physical vapor deposition, low-pressure plasma spraying, cathodic arc deposition, vapor phase aluminide coating, pack cementation, chemical vapor deposition, electroplating, and combinations thereof.
For example, in an electron-beam physical vapor deposition process, sub-segments 24a and 24b are each placed on a rotatable mount in a vacuum chamber containing a target anode derived of the desired coating material. A charged tungsten filament then emits an electron beam that contacts the target anode, thereby ionizing the material of the target anode. The ionized particles then precipitate onto pressure sidewall 40 and suction side wall 42 of vane 26 and onto pressure sidewall 48 and suction side wall 50 of vane 28 to form the desired metallic coatings. Examples of suitable materials for the metallic coatings include aluminum, platinum, MCrAlY alloys, combinations thereof. Examples of suitable average thicknesses for the metallic coatings on vanes 26 and 28 range from about 25 micrometers to about 200 micrometers, with particularly suitable thicknesses ranging from about 50 micrometers to about 100 micrometers.
The above-discussed coating processes may also be used to form thermal barrier coatings on vanes 26 and 28. Suitable materials for the thermal barrier coatings include zirconia-based materials, where the zirconia is desirably modified with a stabilizer to prevent the formation of a monoclinic phase, and pyrochlores. Examples of suitable stabilizers include yttria, gadolinia, calcia, ceria, magnesia, and combinations thereof. Examples of suitable coating thicknesses for the thermal barrier coatings on stator vanes 26 and 28 range from about 25 micrometers to about 1,000 micrometers, with particularly suitable coating thicknesses ranging from about 100 micrometers to about 500 micrometers.
As discussed above, separating sub-segments 24a and 24b allows vanes 26 and 28 to be coated without interference from each other. Thus, when separated, pressure sidewall 40 and suction side wall 42 of vane 26, and pressure sidewall 48 and suction side wall 50 of vane 28, are each outboard surfaces that are directly accessible to the deposited coating materials. Thus, the coatings formed on pressure sidewalls 40 and 48, and on suction side walls 42 and 50, may have substantially uniform thicknesses. This preserves the corrosion, oxidation, and/or thermal resistance of the coatings during the course of operation in a gas turbine engine.
After the coating process is completed, sub-segments are placed together such split edges 72 and 74 of shroud portions 30a and 30b are substantially aligned, and such that split edges 76 and 78 of platform portions 32a and 32b are substantially aligned. Accordingly, during the rejoining process, split edges 72, 74, 76, and 78 function as faying surfaces. Pursuant to steps 20 and 22 of method 10, shroud portions 30a and 30b, and platform portions 32a and 32b are then each rejoined to reform shroud 30 and platform 32, respectively. As discussed above, the rejoining process forms bond lines (not shown in
In one embodiment the rejoining process involves thermal diffusion bonding. In this embodiment, split edges 72 and 74 are compressed together and are subjected to elevated temperatures to interdiffuse the materials of shroud portions 30a and 30b, and split edges 76 and 78 are compressed together and are subjected to elevated temperatures to interdiffuse the materials of platform portions 32a and 32b. This may be performed by compressing sub-segments 24a and 24b together at split edges 72 and 74 and at split edges 76 and 78, and placing the compressed sub-segments 24a/24b in a vacuum oven for a suitable temperature and duration to interdiffuse the materials.
Suitable temperatures and durations for the thermal diffusion bonding include those that sufficiently bond shroud portions 30a and 30b, that sufficiently bond platform portions 32a and 32b, and that also substantially preserve the integrities of the coatings applied to vanes 26 and 28. Examples of suitable temperatures for the thermal diffusion bonding range from about 1040° C. (about 1900° F.) to about 1200° C. (about 2200° F.), with particularly suitable temperatures ranging from about 1090° C. (about 2000° F.) to about 1150° C. (about 2100° F.). Examples of suitable durations for the thermal diffusion bonding include durations up to about one hour, with particularly suitable durations ranging from about 10 minutes to about 30 minutes.
In an alternative embodiment, the rejoining process involves a transient liquid phase bonding. In this embodiment, layers of one or more brazing materials are placed between split edges 72 and 74, and between split edges 76 and 78, and sub-segments 24a and 24b are subjected to elevated temperatures. The elevated temperatures liquefy the brazing materials, thereby allowing the liquefied brazing materials to interdiffuse with the alloys of sub-segments 24a and 24b. Suitable temperatures and durations for the transient liquid phase bonding include those discussed above for the thermal diffusion bonding.
Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.
