Abstract:
An abrasive coating for rotor shafts that interact with cantilevered vanes to form an abradable air seal in a turbine engine. The abrasive coating including a metal bond coat layer on the rotor shaft, and an abrasive top coating bond coat layer for contact with vanes during operation of the rotor shaft, the abrasive coating including a plurality of abrasive grit particles in a matrix. the abrasive grit particles are selected from the group consisting of cubic boron nitride (CBN), zirconia, alumina, silicon carbide, diamond and mixtures thereof.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is related to the following co-pending applications that are filed on even date herewith and are assigned to the same assignee: ABRASIVE ROTOR COATING FOR FORMING A SEAL IN A GAS TURBINE ENGINE, Ser. No. ______, Attorney Docket No. PA0014032U-U73.12-547KL; ROUGH DENSE CERAMIC SEALING SURFACE IN TURBOMACHINES, Ser. No. ______, Attorney Docket No. PA0014043U-U73.12-548KL; THERMAL SPRAY COATING PROCESS FOR COMPRESSOR SHAFTS, Ser. No. ______, Attorney Docket No. PA0014152U-U73.12-549KL; FRIABLE CERAMIC ROTOR SHAFT ABRASIVE COATING, Ser. No. ______, Attorney Docket No. PA0013722U-U73.12-550KL; ABRASIVE ROTOR SHAFT CERAMIC COATING, Ser. No. ______, Attorney Docket No. PA0014199U-U73.12-543KL; LOW DENSITY ABRADABLE COATING WITH FINE POROSITY, Ser. No. ______, Attorney Docket No. PA0013584U-U73.12-541KL; and ABRASIVE CUTTER FORMED BY THERMAL SPRAY AND POST TREATMENT, Ser. No. ______, Attorney Docket No. PA0012340U-U73.12-540KL. The disclosures of these applications are incorporated herein by reference in their entirety. 
     
    
     BACKGROUND 
       [0002]    Gas turbine engines include compressor rotors with a plurality of rotating compressor blades. Minimizing the leakage of air between tips of the compressor blades and a casing of the gas turbine engine increases the efficiency of the gas turbine engine as the leakage of air over the tips of the compressor blades can cause aerodynamic efficiency losses. To minimize this, the gap at tips of the compressor blades is set so small that at certain conditions, the blade tips may rub against and engage an abradable seal on 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. 
         [0003]    Cantilevered vanes that seal against a rotor shaft are also used for elimination of the air leakage in turbine engines. Current cantilevered vane tip sealing requires that the tip gaps need to be set more open than desired for optimum seal in order to prevent rub interactions that can cause rotor shaft spallation, vane damage or rotor shaft burn through caused by thermal runaway events during rubs. Current materials that are primarily ceramics have been shown to lack the durability to prevent spallation and they lack the abradability to prevent vane damage. 
       SUMMARY 
       [0004]    The present invention comprises an abrasive coating on the surface that interacts with the vane tips with a low strength, abrasive composite top layer that contains sharp abrasive grits held in a composite matrix of hexagonal boron nitride (hBN), nickel, chromium, aluminum or NiCrAlY. Examples of sharp abrasive grits are cubic boron nitride (CBN), zirconia, alumina, silicon carbide and diamond. 
         [0005]    The abrasive coating includes a base bond coat layer. The bond coat may be MCr, MCrA, MCrAlY or a refractory modified MCrAlY, where M is nickel, cobalt, iron or mixtures thereof. 
         [0006]    When thermal protection is needed, there is also a layer between the abrasive grit and on the bond coat comprising a ceramic layer that acts as a thermal barrier to protect the rotor shaft. Ceramic layers include zirconia, hafnia, mullite, alumina. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  illustrates a simplified cross-sectional view of a gas turbine engine. 
           [0008]      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. 
           [0009]      FIG. 3  is a cross sectional view taken along the line  3 - 3  of  FIG. 2 , not to scale. 
           [0010]      FIG. 4  is a cross sectional view of another embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]      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 . 
         [0012]    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 . 
         [0013]    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. 
         [0014]    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. 
         [0015]    In operation of turbine engine  10 , incoming airflow F I  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 . 
         [0016]    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. 
         [0017]    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 . 
         [0018]    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. 
         [0019]      FIG. 2  is a cross section along line  22  of  FIG. 1  of a casing  48  which has a rotor shaft  50  inside. For the purpose of illustration, the invention is shown with respect to vanes  26 . The invention can also be used with rotor blades  28 . 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. 
         [0020]      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. 
         [0021]      FIG. 3  shows an embodiment comprising bi-layer coating  60  in which includes metallic bond coat  62  and abradable layer  66 . Metallic bond coat  62  is applied to rotor shaft  50 . Abradable layer  66  is deposited on top of bond coat  62  and is the layer that first encounters vane tip  26 T. In some embodiments, the bond coat  62  can be eliminated because the abradable layer  66  may have a component that provides sufficient bond strength. 
         [0022]    Bond coat  62  is thin, up to 10 mils, more specifically ranging from about 3 mils to about 7 mils (about 76 to about 178 microns). Abradable coating  66  is about the same thickness as bond coat  64 , again ranging from about 3 mils to about 7 mils (about 76 to about 178 microns). 
         [0023]    Bond coat  62  may be formed of MCrAlY, the metal (M) can be nickel, iron, or cobalt, or combinations thereof and the alloying elements are chromium (Cr), aluminum (Al) and yttrium (Y). 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. 
         [0024]    Top abrasive layer  66  is formed from grit particles contained in a low strength abrasive composite. Examples of sharp abrasive grits are CBN, zirconia, alumina, silicon carbide, diamond and mixtures thereof. The matrix holding the abrasive grits is a composite matrix of hBN, Ni, Cr, or MCrAlY. The metal (M) can be nickel, cobalt, iron or mixtures thereof, and the alloying elements are chromium (Cr), aluminum (Al) and yttrium (Y). The grit particles range in size from about 20 microns to about 150 microns. Grit sizes much smaller or larger are less effective as a grit particle. Grit particles in the top layer may also range in size from about 25 to about 75 microns in the composite matrix. 
         [0025]    Because the top abrasive layer  66  includes a metal matrix, bond coat  62  can be eliminated. In some instances, the metallic matrix material described above can be added as a first layer with or without the hBN component. 
         [0026]    The abrasive layer cuts vane tips in a low temperature abrasive manner much like a metal matrix diamond grinding wheel functions. When the grit particles are dulled by excessive use, they are pulled out by the grinding forces and fresh grits are exposed by wear of the matrix. The grits are held in the matrix and cut the vane tips until the grinding forces pull them out to expose fresh grits. 
         [0027]    During slow interactions between grits in the matrix and the vanes, cutting forces are low and little rotor coating wear occurs. When the interaction rates increase, and/or the grit particles no longer cut as well due to increased surface temperatures or dulling, the strength of the matrix is exceeded and the grits fall out. This shedding of overstressed grit exposes the composite matrix to vane tip contact and results in abradable wear. 
         [0028]    Through the balancing of matrix strength and grit content, a balance is achieved between the needs of the engine to round up parts for optimum efficiency, while providing abradable response during high interaction rate events such as take-off, landing and maneuver loading during surges and the like. The strength of the composite ceramic matrix is sufficient to hold and retain sharp grits that cut with low cutting forces. When the grits dull, forces go up and the grits are released, exposing fresh matrix material and grit material. 
         [0029]    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 . 
         [0030]    Optional ceramic layer  64 , shown in  FIG. 4 , may be any of the zirconia based ceramics such as are described in commonly 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 insulated 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. 
         [0031]    The present invention provides for an abrasive layer that interacts with a bare metal surface to abrade the metal and permit effective roundup. In gas turbine engines that are used in flight, the abrasive layer will interact with the bare tip of an airfoil, such as a rotor blade or stator vane. In gas turbine engines that are used on the ground as power stations, the abrasive coating can be on the tip of an airfoil, such as a rotor blade tip or stator vane tip. The abrasive layer abrades the bare metal in all instances, releasing the grit particles when they become dull as noted above. 
         [0032]    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.