Abstract:
A gas turbine engine having an engine casing extending circumferentially about an engine centerline axis; and a compressor section, a combustor section, and a turbine section within said engine casing. At least one of said compressor section and said turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil, wherein a tip of the at least one airfoil is metal having a thin film ceramic coating and the at least one seal member is coated with an abrasive.

Description:
BACKGROUND 
       [0001]    As gas turbine engines evolve to provide better performance, they become hotter, faster and stronger. As a result, the materials used need to be able to function under those increased operating conditions. 
       SUMMARY 
       [0002]    The invention comprises the use of non-abrasive blade tip coatings for use in sliding contact wear between the blade tip coating and an abradable surface. The invention is used in gas turbine engines where the melting point of the blade material is similar to, or lower than, that of the abradable material. 
         [0003]    The coating on the non-abrasive blade tip is a thin ceramic coating that has high hardness, is very smooth and has good mechanical and thermal shock resistance. Thin film ceramic coatings include TiN, TiAlN, Al 2 O 3 , BN, SiCN, TiCN, and TiO. The coating is applied by vapor deposition methods, as conversion coatings, or by slurry application of nano particulate suspensions. These coatings resist adhesion of smeared coating materials during a rub event. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0004]      FIG. 1  illustrates a simplified cross-sectional view of a gas turbine engine. 
           [0005]      FIG. 2  illustrates a simplified cross-sectional view illustrating the relationship of the rotor and vanes taken along the line  2 - 2  of  FIG. 1 , not to scale. 
           [0006]      FIG. 3  is a cross-sectional view taken along the line  3 - 3  of  FIG. 2 , not to scale. 
           [0007]      FIG. 4  illustrates a simplified cross-sectional view illustrating the relationship of the casing or shroud and blades taken along the line  4 - 4  of  FIG. 1 , not to scale. 
           [0008]      FIG. 5  is a cross-sectional view taken along the line  5 - 5  of  FIG. 4 , not to scale. 
       
    
    
     DETAILED DESCRIPTION 
       [0009]    In recent gas turbine engine designs, metal airfoils mate with abradable coatings and have shown evidence of blade metal transfer to the abradable coating resulting in 8 mils of excess clearance in test engines. The high temperatures, blade count and tip speed result in bade tip and coating contact temperature during rub that exceeds historical experience. When this happens, and the blade tip and coating base metals have similar melting points, both materials soften and become prone to adhesive wear mechanisms. Also because blade tip surface area is lower than coating surface area, the tips reach a higher temperature while receiving similar rub energy input compared to the coating. For these reasons, the blade tips become prone to metal transfer to the coating during sliding contact wear. 
         [0010]      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 . 
         [0011]    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  32 , 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 . 
         [0012]    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. 
         [0013]    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. 
         [0014]    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 . 
         [0015]    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 FC, and LPC section  30  (or a portion of it) may be omitted. Alternatively, an additional intermediate spool is included, for example in a three-spool turboprop or turbofan configuration. 
         [0016]    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  42 , driving HPC section  32  of compressor  16  via HPT shaft  42 . 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 . 
         [0017]    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. 
         [0018]    The present invention is intended to be used with airfoils in turbine engines. The term “airfoil” is intended to cover both rotor blades and stator vanes. It is the purpose of this invention to produce non-abrasive blade tip coatings.  FIG. 2  and  FIG. 3  disclose the invention with respect to interaction of a stator vane with a rotor.  FIG. 4  and  FIG. 5  disclose the invention with respect to interaction of a rotor blade with a casing or shroud. The coating of this invention may be used with either or both configurations. 
         [0019]      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 . Abradable coating  60 , is on rotor shaft  50  such that the clearance C between coating  60  and abradable vane tips  26 T of vanes  26  with thin film tip coating  27  has the proper tolerance for operation of the engine, e.g., to serve as a seal to prevent leakage of air (thus increasing 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 25 to 55 mils (635 to 1397 microns) when the engine is cold and 0.000 to 0.035 mils during engine operation depending on the specific operating condition 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  with thin film tip coating  27  that varies with operating conditions, as described herein. Coating  60  is an abradable coating. Coating  27 , described in detail below, is a thin film ceramic coating that has a melting or softening point higher than that of the abradable material  64  of abradable coating  60  and of abradable vane tip  26 T of vane  26 . Coating  27  has high hardness, is very smooth, and has good mechanical properties and thermal shock resistance. In operation, metal transfer to the coating does not take place during sliding contact wear. 
         [0021]    As can be seen from  FIG. 4  and  FIG. 5 , the same concept is used in which coating  70  is provided on the inner diameter surface of casing or shroud  48  and thin film tip coating  29  is provided on tip  28 T of blade  28 . Coating  70  is an abradable coating, Coating  29 , also described in detail below, is also a thin film ceramic coating that has a melting or softening point higher than that of the abradable material and blade tip and blade tip  28 T of blade  28 . Coating  29  has high hardness, is very smooth, and has good mechanical properties and thermal shock resistance. In operation, metal transfer to the coating does not take place during sliding contact wear. 
         [0022]    The invention is suitable for a range of non-abrasive blades and vanes. For aluminum blades and vanes, anodized layers of aluminum oxide are effective. For Ti blades and vanes, anodized layers of titanium dioxide or titanium nitride are effective. Generally, for all materials including Al, Ti, Fe and Ni based alloys, metal oxide, nitride carbide and boride layers are effective. Specifically, aluminum oxide, zirconium oxide, zirconium nitride, chromium oxide, chromium nitride, titanium oxide, titanium nitride, titanium carbo-nitride, titanium aluminum nitride, silicon nitride, silicon carbide, boron nitride boron carbide and tungsten carbide form effective non-abrasive blade tip coatings. 
         [0023]    With respect to boronized surface layers, they are a diffusion case hardening treatment during which boride and di-boride phases are formed in the base metal&#39;s surface. These phases are high melting point very hard phases that will resist wear and metal transfer to the abradable coating. Borides also have low friction and low surface energy, so they will also resist the coating material transfer to the airfoil tips. 
         [0024]    The composition of diffused layer will vary depending on the base material. For example, ferrous materials will form FeB/FeB 2 . Nickel-based alloys will form Ni 4 B 3 /Ni 2 B/Ni 3 B. Cobalt-based alloys form CoB/Co 2 B/Co 3 B. Titanium-based alloys form TiB/TiB 2 . 
         [0025]    The micro-hardness of the diffused layer will vary depending on the base material. For example, FeB/FeB 2  layers will have a micro-hardness in the range of 1600-1900 HV. Other elements, such as Ni, Ti and Co, will produce a different hardness range, some even higher that FeB/FeB 2 . 
         [0026]    Presented below are examples of diffused layers of this invention with the micro-hardness, coefficient of friction values they produce. 
         [0000]    
       
         
               
               
               
               
               
             
               
               
               
               
               
             
           
               
                 TABLE I 
               
               
                   
               
               
                   
                   
                   
                 Micro- 
                   
               
               
                   
                   
                   
                 hardness 
                 Coefficient 
               
               
                 Name 
                 Composition 
                 Color 
                 (HV 0.05 g) 
                 of Friction 
               
               
                   
               
             
             
               
                   
               
             
          
           
               
                 Medikote ™ C 
                 TiN 
                 Gold 
                 2300-2500 
                 0.35 
               
               
                 Medikote ™ C3 
                 CrN 
                 Silver 
                 2000-2200 
                 0.35 
               
               
                 Medikote ™ C5 
                 TiN/TiCN 
                 Bronze/ 
                 2800-3200 
                 0.30 
               
               
                   
                   
                 Gray 
                   
                   
               
               
                 Medikote ™ C6 
                 AlTiN 
                 Violet/ 
                 3000-3400 
                 0.35 
               
               
                   
                   
                 Black 
                   
                   
               
               
                 Medikote ™ C6B 
                 TiAlN 
                 Copper/ 
                 3000-3200 
                 0.4 
               
               
                   
                   
                 Bronze 
                   
                   
               
               
                 Medikote ™ C6JB 
                 AlTiN 
                 Black 
                 3000-3400 
                 0.4 
               
               
                 Medikote ™ C8 
                 ZrN 
                 Pale Gold 
                 2300-2500 
                 0.35 
               
               
                   
               
             
          
         
       
     
         [0027]    The microhardness of the thin film ceramic coating ranges from 2,000 to 3,400 HV 0.05 g. The coefficient of friction of the thin film ceramic coating ranges from 0.35 to 0.40. 
         [0028]    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.