Abstract:
A seal in a gas turbine engine component between an airfoil with a radial outward end and a seal member adjacent it coated with an abrasive layer having a ceramic component in a matrix of a metal alloy with hexagonal BN. The ceramic component is selected from silica, quartz, alumina, zirconia and mixtures thereof and the metal is selected from nickel, cobalt, copper and iron. The ceramic ranges from about 1% to about 10% and the amount of nickel, cobalt, copper or iron will range from about 30% to about 60% by volume, and the balance is hBN.

Description:
BACKGROUND 
       [0001]    Gas turbine engines include compressor rotors having a plurality of rotating compressor blades. Minimizing the leakage of air, such as between tips of rotating blades and a casing of the gas turbine engine, increases the efficiency of the gas turbine engine because the leakage of air over the tips of the blades can cause aerodynamic efficiency losses. To minimize this, the gap at tips of the blades is set small and at certain conditions, the blade tips may rub against and engage an abradable seal at the casing of the gas turbine. The abradability of the seal material prevents damage to the blades while the seal material itself wears to generate an optimized mating surface and thus reduce the leakage of air. 
         [0002]    Cantilevered vanes that seal against a rotor shaft are used for elimination of the air leakage and complex construction of vane inside diameter (ID) shroud, abradable seal and knife edges that are used in present gas turbine engines. Current cantilevered vane tip sealing experiences the difficulty that the tip gaps need to be set more open than desirable to prevent rub interactions that can cause rotor shaft coating spallation, vane damage or rotor shaft burn through due to thermal runaway events during rubs. Current materials have been found to lack the durability to prevent spallation and lack the abradability to prevent vane damage. 
         [0003]    Blade outer seals do not have as many problems as inner seals, but do need to have the ability to resist fine particle erosion and have a suitable wear ratio between the seal and the airfoil. 
         [0004]    It would be an advantage for an abradable coating for rotor that is capable of running against bare vane tips and have a desirable balance of wear between both the vane tips and the coating. The coating should also prevent catastrophic thermal runaway events, coating spallation and damage to the vanes. 
       SUMMARY 
       [0005]    The present invention comprises an abrasive coating forming a seal material on components of gas turbine engines. The present invention comprises an abrasive coating on the surface of the rotor to form a seal with the stator vanes and on the inside of the casing to form a seal with the rotor blades. 
         [0006]    The abrasive coating contains ceramic particles in a composite matrix of hexagonal boron nitride (hBN) in nickel, cobalt, copper, iron or mixtures thereof. The ceramic particles are irregularly flattened shapes that are described as “splats” in the thermal spray field. The ceramic particles may be any ceramic that has a hardness of seven or more on the Mohs Scale for hardness, such as silica, quartz, alumina and zirconia. 
         [0007]    The abrasive coating will often include a base bond coat layer. The bond coat may be MCr, MCrAl., MCrAlY or a refractory modified MCrAlY, where M is nickel, cobalt, iron or mixtures thereof. 
         [0008]    When thermal protection is needed, there is also a layer between the abrasive coating and the bond coat comprising a ceramic layer that acts as a thermal barrier to protect the coated components. Ceramic layers include, for example, zirconia, hafnia, mullite, alumina. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  illustrates a simplified cross-sectional view of a gas turbine engine. 
           [0010]      FIG. 2  illustrates a simplified cross sectional view of a rotor shaft inside a casing illustrating the relationship of the rotor and cantilevered vanes taken along the line  2 - 2  of  FIG. 1 , not to scale. 
           [0011]      FIG. 3  is a cross sectional view taken along the line  3 - 3  of  FIG. 2 , not to scale. 
           [0012]      FIG. 4  is a cross sectional view of another embodiment. 
           [0013]      FIG. 5  is a cross sectional view of yet another embodiment. 
           [0014]      FIG. 6  is a cross sectional view taken along the line  5 - 5  of  FIG. 4 , not to scale. 
       
    
    
     DETAILED DESCRIPTION 
       [0015]      FIG. 1  is a cross-sectional view of gas turbine engine  10 , in a turbofan embodiment. As shown in  FIG. 1 , turbine engine  10  comprises fan  12  positioned in bypass duct  14 , with bypass duct  14  oriented about a turbine core comprising compressor (compressor section)  16 , combustor (or combustors)  18  and turbine (turbine section)  20 , arranged in flow series with upstream inlet  22  and downstream exhaust  24 . 
         [0016]    Compressor  16  comprises stages of compressor vanes  26  and blades  28  arranged in low pressure compressor (LPC) section  30  and high pressure compressor (LPC) section  32 . Turbine  20  comprises stages of turbine vanes  34  and turbine blades  36  arranged in high pressure turbine (HPT) section  38  and low pressure turbine (LPT) section  40 . HPT section  38  is coupled to HPC section  32  via HPT shaft  42 , forming the high pressure spool or high spool. LPT section  40  is coupled to LPC section  30  and fan  12  via LPT shaft  44 , forming the low pressure spool or low spool. HPT shaft  42  and LPT shaft  44  are typically coaxially mounted, with the high and low spools independently rotating about turbine axis (centerline) C L . 
         [0017]    Fan  12  comprises a number of fan airfoils circumferentially arranged around a fan disk or other rotating member, which is coupled (directly or indirectly) to LPC section  30  and driven by LPT shaft  44 . In some embodiments, fan  12  is coupled to the fan spool via geared fan drive mechanism  46 , providing independent fan speed control. 
         [0018]    As shown in  FIG. 1 , fan  12  is forward-mounted and provides thrust by accelerating flow downstream through bypass duct  14 , for example in a high-bypass configuration suitable for commercial and regional jet aircraft operations. Alternatively, fan  12  is an unducted fan or propeller assembly, in either a forward or aft-mounted configuration. In these various embodiments turbine engine  10  comprises any of a high-bypass turbofan, a low-bypass turbofan or a turboprop engine, and the number of spools and the shaft configurations may vary. Also contemplated for use with the present invention are marine and land based turbines that may or may not have a fan or propeller. 
         [0019]    In operation of turbine engine  10 , incoming airflow F 1  enters inlet  22  and divides into core flow F C  and bypass flow F B , downstream of fan  12 . Core flow F C  propagates along the core flowpath through compressor section  16 , combustor  18  and turbine section  20 , and bypass flow F B  propagates along the bypass flowpath through bypass duct  14 . 
         [0020]    LPC section  30  and HPC section  32  of compressor  16  are utilized to compress incoming air for combustor  18 , where fuel is introduced, mixed with air and ignited to produce hot combustion gas. Depending on embodiment, fan  12  also provides some degree of compression (or pre-compression) to core flow F C , and LPC section  30  may be omitted. Alternatively, an additional intermediate spool is included, for example in a three-spool turboprop or turbofan configuration. 
         [0021]    Combustion gas exits combustor  18  and enters HPT section  38  of turbine  20 , encountering turbine vanes  34  and turbine blades  36 . Turbine vanes  34  turn and accelerate the flow, and turbine blades  36  generate lift for conversion to rotational energy via HPT shaft  50 , driving HPC section  32  of compressor  16  via HPT shaft  50 . Partially expanded combustion gas transitions from HPT section  38  to LPT section  40 , driving LPC section  30  and fan  12  via LPT shaft  44 . Exhaust flow exits LPT section  40  and turbine engine  10  via exhaust nozzle  24 . 
         [0022]    The thermodynamic efficiency of turbine engine  10  is tied to the overall pressure ratio, as defined between the delivery pressure at inlet  22  and the compressed air pressure entering combustor  18  from compressor section  16 . In general, a higher pressure ratio offers increased efficiency and improved performance, including greater specific thrust. High pressure ratios also result in increased peak gas path temperatures, higher core pressure and greater flow rates, increasing thermal and mechanical stress on engine components. 
         [0023]      FIG. 2  is a cross section along line  22  of  FIG. 1  of a casing  48  which has a rotor shaft  50  inside. Vanes  26  are attached to casing  48  and the gas path  52  is shown as the space between vanes  26 . Coating  60 , corresponding to the coating of this invention, is on rotor shaft  50  such that the clearance C between coating  60  and vane tips  26 T of vanes  26  has the proper tolerance for operation of the engine, e.g., to serve as a seal to prevent leakage of air (thus reducing efficiency), while not interfering with relative movement of the vanes and rotor shaft. In  FIGS. 2 and 3 , clearance C is expanded for purposes of illustration. In practice, clearance C may be, for example, in a range of about 0.025 inches to 0.055 inches when the engine is cold and 0.000 to 0.035 inches during engine operation, depending on the specific operating conditions and previous rub events that may have occurred. 
         [0024]      FIG. 3  shows the cross section along line  3 - 3  of  FIG. 2 , with casing  48  and vane  26 . Coating  60  is attached to rotor shaft  50 , with a clearance C between coating  60  and vane tip  26 T of vane  26  that varies with operating conditions, as described herein. 
         [0025]      FIG. 3  shows an embodiment comprising bi-layer coating  60  in which includes metallic bond coat  62  and abrasive layer  66 . Metallic bond coat  62  is applied to rotor shaft  50 . Abrasive layer  66  is deposited on top of bond coat  62  and is the layer that first encounters vane tip  26 T. 
         [0026]    Bond coat  62  is thin, up to 10 mils (254 microns), more specifically ranging from about 3 mils to about 7 mils (about 76 to about 178 microns). Abrasive coating  66  may be about the same thickness as bond coat  64 , again ranging from about 3 mils to about 7 mils (about 76 to about 178 microns), while some applications that have larger variation in tip clearance may require a thicker abrasive layer. Abrasive layer  66  may be as thick as 300 mils (7620 microns) in some applications. 
         [0027]    The bond coat may be MCr, MCrAl., MCrAlY or a refractory modified MCrAlY, where M is nickel, cobalt, iron or mixtures thereof. For example, bond coat  62  may be 15-40% Cr 6-15% Al, 0.61 to 1.0%. Y and the balance is cobalt, nickel or iron and combinations thereof. 
         [0028]    Top abrasive layer  66  is a low strength abradable composite matrix of a metal alloy such as Ni, Co, Cu, Al MCrAlY loaded with hexagonal boron nitride (hBN) into which flat ceramic particles have been added by thermal spraying. The amount of Ni to hBN in the abradable matrix ranges from about 30% to about 60% by volume, and more specifically about 40% to about 50% Ni by volume, with the balance being hBN. The Ni alloy, hBN (ahBN) and ceramic may be deposited as a coating by individually feeding the powders to one or more spray torches or by blending the two powders and air plasma spraying (APS). Other spray processes would also be effective, such as combustion flame spray, HVOF, HVAF, LPPS, VPS, HVPS and the like. As part of the coating is a quantity of ceramic that at least partially melts during the spray process to form disc like flat particles, or splat particles. 
         [0029]    The ceramic particles may be any ceramic that has a hardness of seven or more on the Mohs Scale for hardness, such as silica, quartz, alumina and zirconia and that at least partially melts at the spray temperatures. The amount of ceramic in coating  66  ranges from about 1% to about 10% by volume. The amount of metal alloy will range from about 30% to about 60% and more specifically about 40% to about 50% Ni by volume, and the balance of 30% to about 70% by volume of hBN. During the spray application of coating  66 , the porosity of coating  66  is controlled to be less than about 10% and even below 5% to decrease the aerodynamic effect. 
         [0030]    Abrasive layer  66  may also be deposited on an intermediate thermally insulating layer to further protect the rotor shaft from burn through during excessive vane contact.  FIG. 4  shows an embodiment comprising tri-layer coating  60 , which includes intermediate insulating ceramic layer  64  between top abrasive layer  66  and bottom coat layer  62 . 
         [0031]    Optional ceramic layer  64 , shown in  FIG. 4 , may be any of the zirconia based ceramics such as are described in U.S. Pat. Nos. 4,861,618, 5,879,573, 6,102,656 and 6,358,002 which are incorporated by reference herein in their entirety. Zirconia stabilized with 6-8 wt. % yttria is one example of such a ceramic layer  64 . Other examples are zirconia stabilized with ceria, magnesia, mullite, calcia and mixtures thereof. Optional thermally insulating ceramic layer  64  thickness may range from about 7 mils to about 12 mils (about 178 to about 305 microns). In many instances, there is no need for optional thermally insulating ceramic layer  64  because abrasive coating  66  functions to remove material by low temperature abrasion minimizing or eliminating thermal burn through of the rotor in high interaction rate events. 
         [0032]    As can be seen from  FIG. 5  and  FIG. 6 , the same concept is used in which coating  70  is provided on the inner diameter surface of casing or shroud  48 . Coating  70  includes a first metallic bond coat  72  that has been applied to the ID of stator casing  48 . In other embodiments, stator casing  48  includes a shroud that forms a blade air seal. Abrasive layer  76  is formed on metallic bond coating  72  and is the layer that first encounters rotor tip  28 T. 
         [0033]    Coating  66  and  76  has a high abradability during fast and/or deep rubs to prevent catastrophic runaway events and damage to turbine components. During low speed rub interactions when frictional heating is low, the ceramic particles result in the desired wear of airfoil tips. When the interaction rate and rub forces increase for any reason, including local vane material transfer, thermal growth and high interaction rates, rub forces may climb only to a limit. Coating  66  and  76  is designed to have a low enough strength to limit rub forces on the airfoils by abrading at contact pressures of less than about 1,000 psi. In one case, 1,000 psi coating strength relates to about 20 pounds per vane loading of compressor stators. Because the bulk coating must meet the durability requirements of the environment, such as the high G environment of the shaft outside diameter in a cantilevered vane sealing application, the abradable coating  66  and  76  has a strength of greater than about 300 psi. The dual nature of coating  66  and  76  provides high abradability when interaction rates and rub forces increase while also cutting the airfoil when interaction rates are low and the ceramic particles dominate the rub interaction. 
         [0034]    While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.