Patent Document

BACKGROUND 
       [0001]    This disclosure relates to an airfoil for a gas turbine engine. 
         [0002]    Hybrid metal fan blades have been proposed in which a metallic sheath is secured to an aluminum substrate. One example metallic sheath is a titanium structure, which provides for a lightweight airfoil. The sheath is typically secured to a leading edge of the substrate to provide resistance to damage from debris. One approach has been to secure the sheath to the substrate using an adhesive. Unfortunately, in such conventional blades, when a corrosion preventative film adhesive layer was used, it often left a fillet of adhesive at the sheath edge, which inhibited proper urethane coating. 
       SUMMARY 
       [0003]    In one embodiment, an airfoil for a gas turbine engine includes a substrate and a sheath providing an edge. An adhesive secures the sheath to the substrate. The adhesive has a fillet that extends beyond the edge that includes a finished surface. 
         [0004]    In a further embodiment of any of the above, the substrate is a first metal and the sheath is a second metal different than the first metal. 
         [0005]    In a further embodiment of any of the above, the adhesive is configured to provide a barrier between the first and second metals to prevent galvanic corrosion. 
         [0006]    In a further embodiment of any of the above, the adhesive includes a scrim embedded in resin. 
         [0007]    In a further embodiment of any of the above, the scrim is provided beneath the sheath and inboard of the edge. 
         [0008]    In a further embodiment of any of the above, the finished surface includes a scraped contour. 
         [0009]    In a further embodiment of any of the above, the airfoil includes a coating arranged over the substrate and the finished surface. The coating abuts the edge. 
         [0010]    In a further embodiment of any of the above, the airfoil is a fan blade and the sheath provides a leading edge of the airfoil. 
         [0011]    In a further embodiment of any of the above, the sheath includes a flank providing the edge. 
         [0012]    In another embodiment, the airfoil includes a body having first, second, and third surfaces. The first and second surfaces are adjacent to one another and are generally at a right angle to one another. The third surface adjoins the second surface at an obtuse angle and provides a sharp edge configured to scrape a cured adhesive. The first and second surfaces are configured to follow an airfoil sheath contour. 
         [0013]    In a further embodiment of any of the above, a relief aperture adjoins the first and second surfaces to one another and is configured to accommodate a corner of the airfoil sheath contour. 
         [0014]    In another embodiment, a method of manufacturing an airfoil for a gas turbine engine includes the steps of securing a sheath to a substrate with adhesive, curing the adhesive, and mechanically removing a portion of the adhesive extending beyond the sheath. 
         [0015]    In a further embodiment of any of the above, the securing step includes providing a resin-saturated scrim between the sheath and substrate. 
         [0016]    In a further embodiment of any of the above, the curing step includes providing a fillet of adhesive adjoining the sheath and the substrate. 
         [0017]    In a further embodiment of any of the above, the removing step includes scraping the fillet with a tool to provide a finished surface on the adhesive. In a further embodiment of any of the above, the method of manufacturing includes the step of applying a coating over the substrate and the finished surface and adjoining the sheath. The coating provides a fan blade contour along with the sheath. 
         [0018]    In another embodiment, a gas turbine engine includes a fan section. The fan section includes a plurality of fan blades, at least one of said fan blades includes a substrate, a sheath providing an edge, and a cured adhesive that secures the sheath to the substrate. The cured adhesive has a fillet that extends beyond the edge that includes a mechanically worked finished surface. 
         [0019]    In a further embodiment of any of the above, the gas turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor section. 
         [0020]    In a further embodiment of any of the above, the compressor section includes a high pressure compressor section and a low pressure compressor section. The turbine section includes a high pressure turbine section and a low pressure turbine section. The high pressure turbine section is engaged with the high pressure compressor section via a first spool and the low pressure turbine section is engaged with the low pressure compressor section via a second spool. 
         [0021]    In a further embodiment of any of the above, the gas turbine engine includes a geared architecture that engages both the low spool and the fan section. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0023]      FIG. 1  is a schematic, cross-sectional side view of an embodiment of a gas turbine engine. 
           [0024]      FIG. 2  is a perspective view of an embodiment of a fan blade of the engine shown in  FIG. 1 . 
           [0025]      FIG. 3  is a cross-sectional view of the fan blade shown in  FIG. 2  taken along line  3 - 3 . 
           [0026]      FIG. 4  is an enlarged cross-sectional view of the fan blade shown in  FIG. 2  illustrating an adhesive fillet provided between a sheath and a substrate subsequent to curing. 
           [0027]      FIG. 5  is a perspective view of a tool used to remove a portion of the fillet shown in  FIG. 4  to provide a finished surface on the adhesive. 
           [0028]      FIG. 6  is a cross-sectional view of a portion of the fan blade shown in  FIG. 2  with a coating applied over the substrate and the finished surface. 
       
    
    
     DETAILED DESCRIPTION 
       [0029]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
         [0030]    The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0031]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure (or first) compressor section  44  and a low pressure (or first) turbine section  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and high pressure (or second) turbine section  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  supports one or more bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0032]    The core airflow C is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0033]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture  48  is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0034]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned per hour divided by lbf of thrust the engine produces at that minimum point. “Fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
         [0035]    Referring to  FIGS. 2 and 3 , a fan blade  27  of the fan  42  includes a root  31  supporting a platform  34 . An airfoil  35  extends from the platform  34  to a tip  39 . The airfoil  35  includes spaced apart leading and trailing edges  39 ,  41 . Pressure and suction sides  43 ,  45  adjoin the leading and trailing edges  39 ,  41  to provide a fan blade contour  61 . 
         [0036]    The fan blade  27  includes a substrate  53  with an edge  49 . A sheath  47  is secured to the substrate  53  over the edge  49  with adhesive  55 . In one example, the sheath  47  and the substrate  53  are constructed from first and second metals that are different from one another. In one example, the substrate  53  is constructed from an aluminum alloy, and the sheath  47  is constructed from a titanium alloy. It should be understood that other metals or materials may be used. 
         [0037]    The adhesive  55  provides a barrier between the substrate  53  and the sheath  47  to prevent galvanic corrosion. Referring to  FIG. 4 , the adhesive  55  includes a scrim  62  (e.g., a glass scrim) that carries a resin  64 . Examples of the adhesive  55  include a variety of commercially available aerospace-quality metal-bonding adhesives are suitable, including several epoxy- and polyurethane-based adhesive films. In some embodiments, the adhesive  55  is heat-cured via autoclave or other similar means. Examples of suitable bonding agents include type EA9628 epoxy adhesive available from Henkel Corporation, Hysol Division, Bay Point, Calif. and type AF163K epoxy adhesive available from 3M Adhesives, Coatings &amp; Sealers Division, St. Paul, Minn. 
         [0038]    In certain embodiments, such as is shown in  FIG. 3 , the adhesive  55  is a film, which also contributes a minute amount of thickness of blade  27  proximate the sheath  47 . In one example, a layer of adhesive film is about 0.005-0.010 inch (1.2-2.5 mm) thick. Despite the additional thickness, a film-based adhesive allows for generally uniform application, leading to a predictable thickness of airfoil  35  proximate forward airfoil edge  39 . 
         [0039]    Certain adhesives  55 , including the example film-based adhesives above, are compatible with scrim  62 . Scrim  62  provides dielectric separation between airfoil  35  and sheath  47 , preventing galvanic corrosion between the two different metal surfaces of airfoil  35  and sheath  47 . The material forming scrim  62  is often determined by its compatibility with adhesive  55 . One example scrim  62  is a flexible nylon-based layer with a thickness between about 0.005 inch (0.12 mm) and about 0.010 inch (0.25 mm) thick. Other examples of the adhesive  55  and other aspects of the fan blade  27  are set forth in U.S. Patent Application Publication 2011/0211967 to the Applicant, which is incorporated herein by reference in its entirety. 
         [0040]    Returning to  FIG. 3 , the sheath  47  includes first and second flanks  51 ,  91  that are arranged on either side of the edge  49 . The adhesive  55 , when cured, flows beyond the sheath edge and creates a fillet  68  bridging an edge  66  of the sheath  47  and a surface  58  of the substrate  53 . In the area of the fillet  68 , the sheath  47  provides spaced apart interior and exterior surfaces  70 ,  72  adjoined by the edge  66 . A corner  74  is provided at the intersection of the edge  66  and the exterior surface  72 , which may be provided at a generally right angle relative to one another. The scrim  62  is provided beneath the sheath  47  and arranged inboard of the edge  66 . Typically, the fillet  68  is larger than desired and is of variable size, which prevents the desired surface profile of an applied coating  60  over the adhesive  55 , the edge  66  and the surface  58 , as illustrated in  FIGS. 3 and 6 . The coating  60 , which may be urethane, for example, provides the desired fan blade contour  61 . 
         [0041]    To reduce the size of the fillet  68 , a tool  76  is used to mechanically remove a portion of the fillet  68  to provide a mechanically worked finished surface  88 . The adhesive  55  may be cured using a vacuum bag and autoclave, which provides a cured exterior surface having visible attributes such as a relatively smooth texture and/or a glossy or matte surface finish. The mechanically worked surface finish  88 , by way of contrast, will have, for example, striations and/or machining marks left by a tool. The structural characteristics and difference between the cured exterior surface and the mechanically worked surface finish  88  may be appreciated based upon a visual inspection of the part. The mechanically worked finished surface  88  is provided at or below the interior surface  70  to sufficiently expose the edge  66  and provide a desired and consistent bonding surface for the coating  60  between the edge  66  and the surface  58 . 
         [0042]    The tool  76 , which is illustrated in  FIG. 5 , includes first, second, third and fourth surfaces  78 ,  80 ,  82 ,  84 . The first and second surfaces  78 ,  80  are adjacent to one another and arranged at generally a right angle relative to one another. The first and second surfaces  78 ,  80  are respectively configured to follow the exterior surface  72  and the edge  66 . The third surface  82  adjoins the second surface  80  at an obtuse angle. The third surface  82  provides a sharp edge that is configured to scrape the fillet  68  and provide the mechanically worked finished surface  88 . The mechanically worked finished surface  88  includes a scraped contour in the example embodiment. The fourth surface  84  adjoins the third surface  82  and is configured to follow the surface  58  of the substrate  53  without damaging the substrate. Tool surfaces  78  and  84  preferably have rounded edges to preclude damaging the sheath substrate (exterior surface  72 ) or the airfoil substrate (surface  58 ) during the scraping procedure. 
         [0043]    In one example, a relief aperture  86 , which may be a generally circular hole in one example, adjoins the first and second surfaces  78 ,  80  to one another to accommodate the corner  74  of the sheath  47 . Once the mechanically worked finished surface  88  has been provided on the adhesive  55 , the coating  60 , which may be urethane in one example, is applied over the edge  66 , the finished surface  88  and the surface  58  to provide the fan blade contour  61 . 
         [0044]    As a result of the foregoing fan blade embodiment, the problem in conventional blades (i.e., where a corrosion preventative film adhesive layer often left a fillet of adhesive at the sheath edge that inhibited proper urethane coating) has been resolved. 
         [0045]    Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For example, other mechanical methods may be used to remove portions of the fillet  68  to expose the edge  66 . For that reason, the following claims should be studied to determine their true scope and content.

Technology Category: 2