Abstract:
A method of manufacturing a component suitable for use in a gas turbine engine, comprising the steps of forming the component from a substrate having a first surface and a second surface, forming at least one aperture through the substrate from the first surface to the second surface having a first open area, applying a first coating to at least one of the first surface and the second surface adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the first coating, applying a second coating to the first coating adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the second coating, and removing the second coating from the aperture, leaving most or all of the first coating to define a second open area which is smaller than the first open area. In a further aspect, a method of repairing a component suitable for use in a gas turbine engine, the method comprising the steps of removing coatings from the component, repairing any defects in the substrate of the component, and applying coatings as described herein.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application is a divisional of U.S. Ser. No. 11/926,986, filed on Oct. 29, 2007, the entire disclosure of which is incorporated herein by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The technology described herein relates generally to gas turbine engines, and more particularly, to air-cooled components for use in gas turbines and methods of manufacturing and repairing such components. 
         [0003]    A gas turbine engine includes a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited within a combustion chamber for generating hot combustion gases. At least some known combustors include a dome assembly, a cowling, and liners to channel the combustion gases to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. The liners are coupled to the dome assembly with the cowling, and extend downstream from the cowling to define the combustion chamber. 
         [0004]    The operating environment within a gas turbine engine is both thermally and chemically hostile. Significant advances in high temperature alloys have been achieved through the formulation of iron, nickel and cobalt-base superalloys, though components formed from such alloys often cannot withstand long service exposures if located in certain sections of a gas turbine engine, such as the turbine, combustor or augmentor. A common solution is to protect the surfaces of such components with an environmental coating system, such as an aluminide coating or a thermal barrier coating (TBC) system. The latter typically includes an environmentally-resistant bond coat and a thermal barrier coating of ceramic deposited on the bond coat. Bond coats are typically formed from an oxidation-resistant alloy such as MCrAlY where M is iron, cobalt and/or nickel, or from a diffusion aluminide or platinum aluminide that forms an oxidation-resistant intermetallic. 
         [0005]    While thermal barrier coating systems provide significant thermal protection to the underlying component substrate, internal cooling of components such as combustor liners is generally necessary, and may be employed in combination with or in lieu of a thermal barrier coating. Combustor liners of a gas turbine engine often require a complex cooling scheme in which cooling air flows around the combustor and is then discharged into the combustor through carefully configured cooling holes in the combustor liner. The performance of a combustor is directly related to the ability to provide uniform cooling of its surfaces with a limited amount of cooling air. Consequently, processes by which cooling holes and their openings are formed and configured are often critical because the size and shape of each opening determine the amount of air flow exiting the opening and the distribution of the air flow across the surface, and affect the overall flow distribution within the combustor. Other factors, such as local surface temperature of the liner, are also affected by variations in opening size. 
         [0006]    For combustor liners without a thermal barrier coating, cooling holes are typically formed by such conventional drilling techniques as electrical-discharge machining (EDM) and laser machining. However, EDM cannot be used to form cooling holes in a combustor liner having a ceramic TBC since the ceramic is electrically nonconducting, and laser machining is prone to spalling the brittle ceramic TBC by cracking the interface between the substrate and the ceramic. Accordingly, cooling holes have been required to be formed by EDM and/or laser machining prior to applying the TBC system, limiting the thickness of the TBC which can be applied or necessitating a final operation to remove ceramic from the cooling holes in order to reestablish the desired size and shape of the openings. Conventional processes involve protecting cooling holes from TBC deposition or complete removal of applied TBC from the holes to obtain the desired hole geometry. This leaves the underlying metal surface exposed to hostile environmental conditions at the hole locations. 
         [0007]    Current repair methods for air-cooled components such as combustor liners include welding thermal fatigue cracks. The location of openings in the panels, such as cooling or dilution holes, and the use of thermal barrier coatings add additional complexity to the use of welds and patches. In many instances, protective coatings must be removed from an entire panel and/or an entire liner to gain access to the underlying metal itself, then reapplying protective coatings. However, conventional reapplication processes involve protecting cooling holes from TBC deposition or complete removal of applied TBC from the holes to obtain the desired hole geometry. This leaves the underlying metal surface exposed to hostile environmental conditions at the hole locations. In some cases, repair of such panels is not a feasible option, and instead the entire combustor liner is replaced. 
         [0008]    Because conventional designs may rely upon the underlying metal substrate to define the finished hole geometry in the absence of a TBC system applied to the hole surfaces, damage to or repair procedures performed on the holes in the metal substrate may affect the performance of the repaired part. Accordingly, a method is desired for manufacturing air-cooled components such as combustor liners in a manner which is economically and physically feasible, provides enhanced protection to the substrate in the vicinity of the cooling holes, and which yields a satisfactory cooling hole geometry both as-manufactured and as-repaired. 
       BRIEF SUMMARY OF THE INVENTION 
       [0009]    In one aspect, described herein is a component suitable for use in a gas turbine engine. The component includes a substrate defining a surface of the component and has a first surface and a second surface. At least one aperture extends through the substrate from the first surface to the second surface, and has a first open area. The component has a first coating on at least one of the first surface and the second surface adjacent to the at least one aperture. The component also has a second coating overlying the first coating adjacent to the at least one aperture, such that at least a portion of the first coating is exposed adjacent to the at least one aperture. The first coating defines a second open area which is smaller than the first open area. 
         [0010]    In another aspect, described herein is a method of manufacturing a component suitable for use in a gas turbine engine, comprising the steps of forming the component from a substrate having a first surface and a second surface, forming at least one aperture through the substrate from the first surface to the second surface having a first open area, applying a first coating to at least one of the first surface and the second surface adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the first coating, applying a second coating to the first coating adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the second coating, and removing the second coating from the aperture, leaving most or all of the first coating to define a second open area which is smaller than the first open area. 
         [0011]    In a further aspect, described herein is a method of repairing a component suitable for use in a gas turbine engine, the component having a substrate with first and second surfaces and at least one aperture extending through the substrate from the first surface to the second surface, the aperture having a first open area, the method comprising the steps of removing coatings from the component, repairing any defects in the substrate of the component, applying a first coating to at least one of the first surface and the second surface adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the first coating, applying a second coating to the first coating adjacent to the at least one aperture, the aperture remaining at least partially unobstructed by the second coating, and removing the second coating from the aperture, leaving most or all of the first coating to define a second open area which is smaller than the first open area. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]    The accompanying drawings illustrate several embodiments of the technology described herein, wherein: 
           [0013]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine; 
           [0014]      FIG. 2  is a schematic cross-sectional view of an exemplary combustor assembly that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0015]      FIG. 3  is an enlarged perspective view of a portion of an exemplary combustor liner that may be used with the combustor assembly shown in  FIG. 2 ; 
           [0016]      FIG. 4  is an enlarged partial cross-sectional view of the combustor liner shown in  FIG. 3  before a coating application; and 
           [0017]      FIG. 5  is an enlarged partial cross-sectional view of the combustor liner shown in  FIG. 4  after a coating application; and 
           [0018]      FIG. 6  is an enlarged partial cross-sectional view of the combustor liner shown in  FIG. 5  after removing some coating material; and 
           [0019]      FIG. 7  is a flowchart illustrating steps associated with an exemplary manufacturing method; and 
           [0020]      FIG. 8  is a flowchart illustrating steps associated with an exemplary repair method. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0021]    The present invention is generally applicable to air-cooled components, and particularly those that are protected from a thermally and chemically hostile environment by a thermal barrier coating system. Notable examples of such components include the high and low pressure turbine nozzles and blades, shrouds, combustor liners and augmentor hardware of gas turbine engines. The advantages of this invention are particularly applicable to gas turbine engine components that employ internal cooling and a thermal barrier coating to maintain the service temperature of the component at an acceptable level while operating in a thermally hostile environment. 
         [0022]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine  10 . Engine  10  includes a low pressure compressor  12 , a high pressure compressor  14 , and a combustor assembly  16 . Engine  10  also includes a high pressure turbine  18 , and a low pressure turbine  20  arranged in a serial, axial flow relationship. Compressor  12  and turbine  20  are coupled by a first shaft  21 , and compressor  14  and turbine  18  are coupled by a second shaft  22 . In the exemplary embodiment, gas turbine engine  10  is a CFM-56 engine commercially available from CFM International, Inc., Cincinnati, Ohio. In another embodiment, gas turbine engine  10  is a CF-34 engine commercially available from GE&#39;s Aviation business, Cincinnati, Ohio. 
         [0023]    In operation, air flows through low pressure compressor  12  and compressed air is supplied from low pressure compressor  12  to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Airflow from combustor  16  drives turbines  18  and  20  and exits gas turbine engine  10  through a nozzle (not numbered). 
         [0024]      FIG. 2  is a schematic cross-sectional view of an exemplary combustor  16  that may be used with gas turbine engine  10  (shown in  FIG. 1 ). Combustor  16  includes an outer liner  52  and an inner liner  54  disposed between an outer combustor casing  56  and an inner combustor casing  58 . Outer and inner liners  52  and  54  are spaced radially from each other such that a combustion chamber  60  is defined therebetween. Outer liner  52  and outer casing  56  form an outer passage  62  therebetween, and inner liner  54  and inner casing  58  form an inner passage  64  therebetween. A cowl assembly  66  is coupled to the upstream ends of outer and inner liners  52  and  54 , respectively. An annular opening  68  formed in cowl assembly  66  enables compressed air entering combustor  16  through a diffuse opening in a direction generally indicated by arrow A. The compressed air flows through annular opening  68  to support combustion and to facilitate cooling liners  52  and  54 . 
         [0025]    An annular dome plate  70  extends between, and is coupled to, outer and inner liners  52  and  54  near their upstream ends. A plurality of circumferentially spaced swirler assemblies  72  are coupled to dome plate  70 . Each swirler assembly  72  receives compressed air from opening  68  and fuel from a corresponding fuel injector  74 . Fuel and air are swirled and mixed together by swirler assemblies  72 , and the resulting fuel/air mixture is discharged into combustion chamber  60 . Combustor  16  includes a longitudinal axis  75  which extends from a forward end  76  to an aft end  78  of combustor  16 . In the exemplary embodiment, combustor  16  is a single annular combustor. Alternatively, combustor  16  may be any other combustor, including, but not limited to a double annular combustor. 
         [0026]    In the exemplary embodiment, outer and inner liners  52  and  54  each include a plurality of overlapped panels  80 . More specifically, in the exemplary embodiment, outer liner  52  includes five panels  80  and inner liner  54  includes four panels  80 . In an alternative embodiment, both outer and inner liner  52  and  54  may each include any number of panels  80 . Panels  80  define combustion chamber  60  within combustor  16 . Specifically, in the exemplary embodiment, a pair of first panels  82 , positioned upstream, define a primary combustion zone  84 , a pair of second panels  86 , positioned downstream from first panels  82 , define an intermediate combustion zone  88 , and a pair of third panels  90 , positioned downstream (direction B in  FIG. 3 ) from second panels  86 , and a pair of fourth panels  92 , positioned downstream from third panels  90 , define a downstream dilution combustion zone  94 . 
         [0027]    Combustor liners may include dilution holes to provide air into the combustion environment with the combustor, such as to alter the temperature distribution or combustion characteristics. Dilution air is introduced primarily into combustor chamber  60  through a plurality of circumferentially spaced dilution holes  96  that extend through either or both of outer and inner liners  52  and  54 . In the exemplary embodiment, dilution holes  96  are each substantially circular. Dilution holes may be adapted (sized, shaped, and/or arranged) as needed to accomplish the durability and performance objectives of the particular component and the particular product application. 
         [0028]      FIG. 3  illustrates an exemplary combustor liner  52  that may be used with combustor  16 . Liner  52  also includes a plurality of cooling holes  160  formed in the third panel  90  that facilitate cooling liner  52 . Although, only one group of cooling holes  160  is illustrated in the third panel  90 , it should be understood that the group of cooling holes  160  are spaced circumferentially about the third panel  90 . It should be appreciated that each group of cooling holes  160  is positioned corresponding hot spots to facilitate channeling cooling fluid onto the corresponding hot spot. Third panel  90  includes any number of cooling holes  160  that facilitates cooling of liner  52 . 
         [0029]    During operation of gas turbine engine  10 , an inner surface  33  of liner  52  becomes hot and requires cooling. Consequently, in the exemplary embodiment, cooling features such as cooling holes  160  are positioned in liner  52  to facilitate channeling cooling fluid onto hot spots of liner  52 . More specifically, cooling holes  160  channel cooling fluid from outer passage  62  and/or inner passage  64  to the combustion chamber  60 , thus providing a layer of cooling fluid to inner surface  33 . It should be appreciated that other embodiments may use any configuration of cooling holes  160  that enables cooling holes  160  to function as described herein. Similarly, holes  160  could be in liner  54  to cool its outer surface. 
         [0030]    During operation, as atomized fuel is injecting into combustion chamber  60  and ignited, heat is generated within combustion chamber  60 . Although air enters combustion chamber  60  through cooling features  160  and forms a thin protective boundary of air along combustor liner surface  33 , a variation in exposure of combustor liner surfaces to high temperatures may induce thermal stresses into panels  80 . As a result of continued exposure to thermal stresses, over time, panels  80  may become deteriorated. 
         [0031]      FIG. 4  is an enlarged partial cross-sectional view of a portion of the combustor liner  52  to illustrate the relationship between the cooling hole  160  and the liner surface  33 , as well as the axis  220  of the hole  160 . 
         [0032]    Referring now to  FIGS. 5 and 6 , a layer  210  of thermal barrier material is applied to the combustor liner  52  shown in  FIG. 4  on combustor liner surface  33 . Thermal barrier material further insulates combustor liner surface  33  from high temperature combustion gases. Layer  210  includes an inner layer  212 , such as a bond coat layer, and an outer layer  214 , such as a thermal barrier layer. 
         [0033]    The exemplary methods will be described in terms of an air-cooled component, such as a combustor liner  52 , whose metallic substrate  33  is protected by a thermal barrier coating system composed of a bond coat  212  formed on the substrate (inner surface  33 ), and a ceramic layer  214  adhered to the surface  33  with the bond coat  212 . Bond coat  212  and ceramic layer  214  may each be a single layer of material, or formed of two or more layers (i.e., multi-layer) of appropriate materials. As is the situation with high temperature components of a gas turbine engine, the surface  33  may be an iron, nickel or cobalt-base superalloy. The bond coat  212  is preferably an oxidation-resistant composition, such as a diffusion aluminide or MCrAlY, that forms an alumina (Al 2 O 3 ) layer or scale (not shown) on its surface during exposure to elevated temperatures. The alumina scale protects the underlying superalloy surface  33  from oxidation and provides a surface to which the ceramic layer  214  more tenaciously adheres. 
         [0034]    The ceramic layer  214  can be deposited by air plasma spraying (APS), low pressure plasma spraying (LPPS), or physical vapor deposition (PVD) techniques such as electron beam physical vapor deposition (EBPVD), the latter of which yields a strain-tolerant columnar grain structure. An exemplary material for the ceramic layer  214  is zirconia partially stabilized with yttria (yttria-stabilized zirconia, or YSZ), though zirconia fully stabilized with yttria could be used, as well as zirconia stabilized by other oxides, such as magnesia (MgO), calcia (CaO), ceria (CeO 2 ) or scandia (Sc 2 O 3 ). 
         [0035]    The method of this invention entails producing a cooling hole  160  (shown in  FIGS. 4-6 ) which can project through the ceramic layer  214 , bond coat  212  and surface  33  via an opening  162 , to achieve a configuration for the cooling hole  160  and opening  162  that provides an appropriately metered distribution of cooling air across the external surface of the component, such as liner  52 . As shown in  FIG. 5 , the cooling hole opening  162  as initially coated forms a small opening (having axis  230 ) aimed at a steep angle (angle  13 ) to the surface. As shown in  FIG. 6 , after removing the portion of the ceramic layer  214  that is in alignment with the hole, the opening  162  is at a relatively shallow angle to the surface  33  such that the cooling air flowing through the opening  162  can be laid down as an effective film over the component surface during operation. 
         [0036]      FIGS. 7 and 8  illustrate in flow diagram form the exemplary methods described in greater detail herein. While both methods share some common steps, the method  200  is particularly suited for manufacturing of new air-cooled components while the method  300  is particularly suited for repair and restoration of air-cooled components during their service life. 
         [0037]    As shown in  FIG. 4 , a first step of this exemplary method is to form a hole  160  through the liner  52 . A second step is then to apply as shown in  FIG. 5  the bond coat  212  and ceramic layer  214  to the surface  33 . Due to coating buildup at the edges of the hole  160 , the resulting hole opening  162  is smaller in cross-sectional diameter than the cooling hole  160  required for the liner  52 , but is not completely obstructed such that the location of the hole and at least a portion of its cross-section remain substantially free of obstruction. For example, for a cooling hole  160  having a diameter of about 0.035 inch (about 0.9 mm) to about 0.040 inch (about 1.0 mm), the opening  162  after coating preferably has a diameter of about 0.020 inch (about 0.5 mm), or roughly half that intended for the cooling hole  160 , such that a “witness hole” remains visible and accessible through the coatings. Suitable techniques for forming the hole  160  include EDM, though it is foreseeable that the hole  160  could be formed by such other methods as casting, laser, or drilling with an abrasive water jet. As a result of the drilling operation, the hole  160  has a substantially uniform circular cross-section, and forms a non-normal angle (angle a) to the surface  33 . 
         [0038]    Once the hole  160  is formed, and the bond coat  212  and ceramic layer  214  are applied, the component (liner  52 ) is processed through a carefully controlled operation that uses a pressurized fluid stream targeted at the hole  160 , such as from the uncoated side of the liner  52 , to produce the cooling hole  160  and opening  162  shown in  FIG. 5 . Various fluids could be used, such as air or water, containing a media such as glass beads or an abrasive grit to provide an abrasive action on coating materials overlying the hole  160 . 
         [0039]    An operation as described herein has been found to provide sufficient energy to enlarge the opening  162  to the size desired as well as the angle desired by removing the ceramic TBC layer but not the bond coat layer or underlying parent material such as the metal substrate. Therefore, while the operation removes the ceramic layer  214  most or all of the underlying bond coat  212  remains on the surface of the opening adjacent to the cooling hole  160 , such that the bond coat layer provides protection for the edges of the liner in the vicinity of the cooling hole both during manufacture and in service. Because the operation uses mechanical energy rather than heat energy, it does not damage or spall the bond coat  212  or ceramic layer  214  surrounding the hole  160  and forming the edges of the resulting hole opening  162 . 
         [0040]    The method is capable of appropriately sizing and shaping cooling holes and openings through a ceramic thermal barrier coating (TBC) and its underlying substrate. The abrasive fluid stream also serves to finish the hole and its opening, including the desired size and shape of the hole and opening, without removing or damaging the ceramic surrounding the cooling hole and opening. 
         [0041]    If a field returned engine, such as engine  10 , indicates that combustor liner  52  includes at least one deteriorated panel  80 , a variety of repair methods may be employed to restore combustor liner  52  to serviceable condition. These repair methods may include replacement of the entire liner, a complete panel, and/or a portion or segment of a liner panel, as well as repair of cracks such as by welding them closed. 
         [0042]    During a repair operation, all dirt, foreign material, and coatings are normally removed from a component such as a combustor liner to permit a detailed inspection of the component. Any defects in the substrate, such as cracks, are then repaired using suitable and approved methods such as welding, brazing, or replacement of discrete sections of the component. Holes such as cooling holes may be redrilled and/or repaired as needed to restore them to the appropriate size, shape, and pattern. 
         [0043]    Once the surfaces of the component have been suitably repaired, protective thermal barrier coatings may be applied to component surfaces utilizing the exemplary methods described above. Because the finished opening dimensions are carefully controlled and are defined by a removable and replaceable coating system as described herein, it is possible to perform and repeat the repair process while maintaining finished cooling hole dimensions within specifications. 
         [0044]    Because components such as deteriorated liners are repaired using the method described herein, utilizing readily available coating techniques, combustors may be returned to service using a repair process that facilitates improved savings in comparison to removing and replacing entire combustor liners or large patches or complete panels. 
         [0045]    Although the apparatus and methods described herein are described in the context of cooling holes in a combustor liner of a gas turbine engine, it is understood that the apparatus and methods are not limited to gas turbine engines, combustor liners, or cooling holes. Likewise, the gas turbine engine and combustor liner components illustrated are not limited to the specific embodiments described herein, but rather, components of both the gas turbine engine and the combustor liner can be utilized independently and separately from other components described herein. 
         [0046]    While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.