Patent Publication Number: US-2016230557-A1

Title: Hot section repair of metallic coatings

Description:
BACKGROUND 
     In an aircraft environment, an engine that is used to provide thrust to the aircraft may include a turbine that is used to extract energy provided by a combustor for driving a compressor. The turbine typically includes airfoils. The airfoils may have thin walls that are subjected to wear as the engine is operated over a temperature cycle/range. 
     These airfoils may need to be repaired. Traditionally, a stripping technique is used to remove the entirety of a bondcoat and any diffusion zones/material between the bondcoat and the metal base alloy of the airfoil. Then, a low pressure plasma spraying (LPPS) technique is used to repair/restore the coating. However, application of the stripping technique and LPPS to thin walled airfoils can cause one or both of two conditions: (1) a removal of a portion of the metal base alloy, or (2) an excess of coating (e.g., coating in an amount greater than a threshold) to be deposited onto the airfoil, such that the airfoil has a dimension/thickness that exceeds a tolerable threshold/limit. Either condition may lead to having to scrap, restrip/recoat, or throw away the airfoil. 
     BRIEF SUMMARY 
     The following presents a simplified summary in order to provide a basic understanding of some aspects of the disclosure. The summary is not an extensive overview of the disclosure. It is neither intended to identify key or critical elements of the disclosure nor to delineate the scope of the disclosure. The following summary merely presents some concepts of the disclosure in a simplified form as a prelude to the description below. 
     Aspects of the disclosure are directed to a method for servicing an airfoil of an aircraft engine, comprising: removing a first portion of a first bondcoat layer from the airfoil while leaving a second portion of the first bondcoat layer intact on the airfoil, and applying a second bondcoat layer to the airfoil using a coating technique subsequent to the removal of the first portion of the first bondcoat layer. In some embodiments, the method further comprises removing a ceramic layer from the airfoil. In some embodiments, the ceramic layer is removed from the airfoil prior to the removal of the first portion of the first bondcoat layer from the airfoil. In some embodiments, the ceramic layer is removed from the airfoil based on an application of a stripping or blasting technique. In some embodiments, the first portion of the first bondcoat layer is removed from the airfoil based on an application of an acid. In some embodiments, the method further comprises removing the application of the acid prior to applying the second bondcoat layer to the airfoil. In some embodiments, the method is applied as part of a scheduled maintenance activity associated with the engine. In some embodiments, the method is applied as part of an unscheduled maintenance activity associated with the engine. In some embodiments, the airfoil comprises a blade associated with a turbine of the engine. 
     Aspects of the disclosure are directed to an airfoil of an aircraft engine, comprising a first bondcoat layer, and a second bondcoat layer that is applied to the airfoil via a coating technique subsequent to a removal of a first portion of the first bondcoat layer from the airfoil. In some embodiments, the first portion of the first bondcoat layer is removed based on an application of an acid. In some embodiments, the acid is removed prior to the application of the second bondcoat layer to the airfoil. In some embodiments, the second bondcoat layer is applied as part of a scheduled maintenance activity associated with the engine. In some embodiments, the second bondcoat layer is applied as part of an unscheduled maintenance activity associated with the engine. In some embodiments, the airfoil comprises a blade associated with a turbine of the engine. 
     Aspects of the disclosure are directed to a method for servicing hardware associated with an aircraft engine, comprising: removing a first portion of a first layer from the hardware while leaving a second portion of the first layer intact on the hardware; and applying a second layer to the hardware using a coating technique subsequent to the removal of the first portion of the first layer. In some embodiments, the hardware comprises at least one of a turbine blade, a vane, a seal, a combustor float wall panel, or a nozzle. In some embodiments, at least one of the first layer or the second layer comprises a bondcoat. In some embodiments, at least one of the first layer or the second layer comprises a metallic coating. In some embodiments, the coating technique comprises a cathodic arc technique. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present disclosure is illustrated by way of example and not limited in the accompanying figures in which like reference numerals indicate similar elements. 
         FIG. 1  is a side cutaway illustration of a geared turbine engine. 
         FIG. 2  illustrates an exemplary cathodic arc apparatus. 
         FIG. 3A  illustrates an exemplary airfoil. 
         FIG. 3B  illustrates a stack-up of materials of the airfoil of  FIG. 3A . 
         FIG. 4  illustrates a flowchart of an exemplary method. 
     
    
    
     DETAILED DESCRIPTION 
     It is noted that various connections are set forth between elements in the following description and in the drawings (the contents of which are included in this disclosure by way of reference). It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect. A coupling between two or more entities may refer to a direct connection or an indirect connection. An indirect connection may incorporate one or more intervening entities. 
     In accordance with various aspects of the disclosure, apparatuses, systems and methods are described for manufacturing or repairing one or more components, such as an airfoil of a turbine. In some embodiments, engine parts/components may be stripped of a ceramic and bondcoat (e.g., a low pressure plasma spraying (LPPS) metallic bondcoat) may be partially stripped to remove an oxidized or depleted coating. In some embodiments, a cathodic arc deposition technique may be used to apply a thin coating with correct bondcoat chemistry on top of the partially stripped bondcoat. Aspects of the disclosure may be applied in connection with a metallic coating, potentially in lieu of use or application in connection with a bondcoat. 
       FIG. 1  is a side cutaway illustration of a geared turbine engine  10 . This turbine engine  10  extends along an axial centerline  12  between an upstream airflow inlet  14  and a downstream airflow exhaust  16 . The turbine engine  10  includes a fan section  18 , a compressor section  19 , a combustor section  20  and a turbine section  21 . The compressor section  19  includes a low pressure compressor (LPC) section  19 A and a high pressure compressor (HPC) section  19 B. The turbine section  21  includes a high pressure turbine (HPT) section  21 A and a low pressure turbine (LPT) section  21 B. 
     The engine sections  18 - 21  are arranged sequentially along the centerline  12  within an engine housing  22 . Each of the engine sections  18 - 19 B,  21 A and  21 B includes a respective rotor  24 - 28 . Each of these rotors  24 - 28  includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s). 
     The fan rotor  24  is connected to a gear train  30 , for example, through a fan shaft  32 . The gear train  30  and the LPC rotor  25  are connected to and driven by the LPT rotor  28  through a low speed shaft  33 . The HPC rotor  26  is connected to and driven by the HPT rotor  27  through a high speed shaft  34 . The shafts  32 - 34  are rotatably supported by a plurality of bearings  36 ; e.g., rolling element and/or thrust bearings. Each of these bearings  36  is connected to the engine housing  22  by at least one stationary structure such as, for example, an annular support strut. 
     During operation, air enters the turbine engine  10  through the airflow inlet  14 , and is directed through the fan section  18  and into a core gas path  38  and a bypass gas path  40 . The air within the core gas path  38  may be referred to as “core air”. The air within the bypass gas path  40  may be referred to as “bypass air”. The core air is directed through the engine sections  19 - 21 , and exits the turbine engine  10  through the airflow exhaust  16  to provide forward engine thrust. Within the combustor section  20 , fuel is injected into a combustion chamber  42  and mixed with compressed core air. This fuel-core air mixture is ignited to power the turbine engine  10 . The bypass air is directed through the bypass gas path  40  and out of the turbine engine  10  through a bypass nozzle  44  to provide additional forward engine thrust. This additional forward engine thrust may account for a majority (e.g., more than 70 percent) of total engine thrust. Alternatively, at least some of the bypass air may be directed out of the turbine engine  10  through a thrust reverser to provide reverse engine thrust. 
     The engine  10  is illustrative. Aspects of the disclosure may be applied in connection with other engine types or configurations. For example, aspects of the disclosure may be applied in connection with, e.g., aerospace and non-aerospace turbine engines, such as for example locomotive engines, tank engines, small industrial gas turbines, etc. 
     One or more portions of the engine  10 , such as the compressor section  19  or the turbine section  21 , may include one or more airfoils. An illustrative example of an airfoil  300  is shown in  FIG. 3A . 
     The airfoil  300  may be composed of a first section  302 , referred to as a blade. A second section  304  may be configured to attach the airfoil  300  to a rotor of a turbine (e.g., the turbine section  21  of  FIG. 1 ). 
     Referring to  FIG. 3B , a stack-up of materials associated with at least a portion of the airfoil  300  (e.g., the blade  302 ) is shown. A metal base layer  352  is coupled to a bondcoat layer  362 , which in turn is coupled to a ceramic layer  372 . 
     In some embodiments the bondcoat layer  362  may be approximately 0.001 inches to 0.005 inches thick and the ceramic layer  372  may be approximately 0.0075 inches to 0.125 inches thick. Other values may be used in some embodiments. 
       FIG. 3B  is representative of one potential stack-up of materials that may be used. The bondcoat layer  362  may be used to provide oxidation resistance. In some embodiments, a thermal barrier coating (TBC) may be included to provide thermal protection. In some embodiments, a stack-up may include a base alloy with a metallic layer, which may be representative of stand-alone metallic coatings and under platform coatings. A metallic coating may be used for oxidation/corrosion resistance. 
       FIG. 2  is a schematic drawing of a cathodic arc apparatus  200 . The cathodic arc process occurs within an inner chamber  212 , which is surrounded by an outer vacuum chamber  210 . A vacuum chamber  210  is provided with fluid cooling via a coolant supply  213 . The vacuum chamber  210  has provisions for evacuation to provide a relatively high vacuum environment. A cathode  214  is located relatively centrally in the inner chamber  212 . One or more power supplies  222  cause arcing to occur between inner chamber walls  215  and a cylindrical surface  216  of the cathode  214 . Assembly  218  may contain an assembly of magnets (not shown), which can be manipulated to influence the arc position and motion. Assembly  218  may provide conductive cooling to the cathode  214 . The components  220  (e.g., the airfoil  300 ) to be coated are located around the cathode  214  with provisions (not shown) to rotate and otherwise manipulate the components  220  so as to promote the formation of a uniform coating on the desired surfaces. In some circumstances the manipulation may be employed to put a coating of controlled, but varying, thickness on a component  220 . 
     The apparatus  200  may include a substrate electrical bias source  224 , which may be used to provide an electrical bias to a substrate as part of a cathodic arc deposition technique using the apparatus  200 . The apparatus  200  may include a mechanical rough vacuum pump  226  and/or a high volume vacuum pump  228 . These pumps  226  and  228  may be associated with the vacuum chamber  210 , and their role/function would be understood by one of skill in the art. 
     The apparatus  200  is illustrative. Other types/configurations of a cathodic arc apparatus may be used in accordance with aspects of the disclosure. 
     Referring to  FIG. 4 , a flow chart of a method  400  is shown. The method  400  may be executed by, or applied in connection with, one or more systems, components, or devices, such as the apparatus  200  of  FIG. 2  and/or the airfoil  300  of  FIGS. 3A-3B . 
     In block  402 , a maintenance activity is initiated. The maintenance activity of block  402  may be a scheduled or unscheduled maintenance activity, and may be associated with an event where an engine (e.g., the engine  10 ) is disassembled or subject to service. 
     In block  404 , an airfoil (e.g., airfoil  300 ) may be removed from the engine. 
     In block  406 , a ceramic layer (e.g., ceramic layer  372 ) may be removed from the airfoil. The removal of the ceramic layer in block  406  may be based on an application of a stripping or blasting technique. In another aspect, the airfoil  300  is obtained from an inventory location and the process starts at step  406 . 
     In block  408 , an acid may be applied to the airfoil, or a portion of the airfoil (e.g., the blade  302 ), to partially remove a first bondcoat layer (e.g., bondcoat layer  362 ). The amount of removal of the first bondcoat layer may be based on an exposure time of the airfoil to the acid. As a result of block  408 , a first portion of the first bondcoat layer may be removed, leaving a second portion of the first bondcoat layer intact. 
     In block  410 , the acid may be removed from the airfoil. 
     In block  412 , a second bondcoat layer may be applied to the airfoil using, e.g., a cathodic arc technique. In some embodiments, a ceramic layer may be reapplied as part of block  412 . 
     The method  400  is illustrative. In some embodiments, one or more of the blocks may be optional. In some embodiments, the blocks may execute in an order or sequence that is different from what is shown. In some embodiments, additional blocks or operations that are not shown may be included. 
     Aspects of the disclosure may be applied in connection with various types of hardware. For example, aspects of the disclosure may be used to service one or more of a turbine blade, a vane, a seal, a combustor float wall panel, or a nozzle. 
     While some of the examples described above related to the use of a cathodic arc techniques, aspects of the disclosure may be applied in connection with various types of coating techniques. Such techniques may be used to apply a controlled thin coating. 
     Some of the examples described above related to a manufacture or repair of hardware. The servicing may apply to new or “engine run” hardware/parts. In some instances, servicing may include a stripping of one or more coatings during original manufacture to rework an improperly applied coating. 
     Technical effects and benefits of this disclosure include an ability to save engine hardware from undergoing a traditional strip and coat process which would take wall thickness away from the casting and result in scrapped out hardware. Use of a cathodic arc repair technique may be used to restore coating chemistry and coating thickness (of a partially stripped coating) without removing wall thickness from the hardware. Accordingly, a lifetime of a component (e.g., an airfoil, such as a vane or a blade) that is subject to such techniques may be extended. 
     Aspects of the disclosure have been described in terms of illustrative embodiments thereof. Numerous other embodiments, modifications, and variations within the scope and spirit of the appended claims will occur to persons of ordinary skill in the art from a review of this disclosure. For example, one of ordinary skill in the art will appreciate that the steps described in conjunction with the illustrative figures may be performed in other than the recited order, and that one or more steps illustrated may be optional in accordance with aspects of the disclosure. One or more features described in connection with a first embodiment may be combined with one or more features of one or more additional embodiments.