Patent Document

CROSS-REFERENCE TO RELATED APPLICATIONS 
     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. 12/910,989; ROUGH DENSE CERAMIC SEALING SURFACE IN TURBOMACHINES, Ser. No. 12/910,973; THERMAL SPRAY COATING PROCESS FOR COMPRESSOR SHAFTS, Ser. No. 12/910,994; FRIABLE CERAMIC ROTOR SHAFT ABRASIVE COATING, Ser. No. 12/910,966; ABRASIVE ROTOR SHAFT CERAMIC COATING, Ser. No. 12/910,960; LOW DENSITY ABRADABLE COATING WITH FINE POROSITY, Ser. No. 12/910,982; and SELF DRESSING, MILDLY ABRASIVE COATING FOR CLEARANCE CONTROL, Ser. No. 12/910,954. The disclosures of these applications are incorporated herein by reference in their entirety. 
     BACKGROUND 
     Gas turbine engines include compressor rotors including 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. 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. 
     Abradable seals have also been used in turbines to reduce the gap between a rotor and a vane. Thermally sprayed abradable seals have been used in gas turbine engines since the late 1960s. The seals have been made as coatings from composite materials that derive their abradability from the use of low shear strength materials or from a porous, friable coating. 
     The high conductivity of conventional alumina rotor coatings causes thermal runaway events when rub occurs between vanes and the rotor shaft. The runaway event is caused by heat generation during rub that raises the temperature of both the vane tips and the rotor shaft, especially when the rub contact is limited to only a portion of the rotor&#39;s circumference. The heat generated causes expansion of the parts, which increases the rub forces, leading to more heat and then more expansion. The cycle becomes self propagating and has resulted in rotor shaft burn through. 
     In the past, cantilevered vane rubs have been typically limited to less than 2 mils (50.4 microns) and have less than full circumference contact due to the risks of high rub forces, coating spallation or a thermal runaway event where the heat from the rub causes thermal expansion of the rotor. The rotor, when heated sufficiently, can grow out to interfere with the vanes. The result can be a burn through causing holes in the rotating shaft, which can cause subsequent unscheduled engine removal. 
     SUMMARY 
     The present invention comprises a gas turbine engine component and the method of making the same. The component includes an airfoil with a radial outword end and a radial inward end that is to be used with a seal member adjacent to the radial inward end of the airfoil. The seal member is coated with a ceramic layer which is then processed with a laser to have a laser engraved surface in which the top of the surface has less than about 5% of the surface area of the base of the ceramic layer. 
     The ceramic layer has a hardness of at least 7 on the Mohs mineral hardness scale. Examples of ceramics that form this ceramic layer are quartz, cubic zirconia, corundrum and diamond. The thickness should range from about 50 microns to about 500 microns. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a simplified cross-sectional view of a standard gas turbine engine. 
         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. 
         FIG. 3  is a cross sectional view taken along the line  3 - 3  of  FIG. 2 , not to scale. 
         FIG. 4  is a cross sectional view taken along line  4 - 4  of  FIG. 1  of one embodiment of the invention. 
         FIG. 5  is a cross sectional view taken along the line  5 - 5  of  FIG. 4 , not to scale. 
         FIGS. 6 and 7  are photographs of embodiments of this invention. 
     
    
    
     DETAILED DESCRIPTION 
       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 . 
     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 . 
     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. 
     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. 
     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 . 
     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. 
     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 . 
     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. 
     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.  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 stator casing or shroud. The coating of this invention may be used with either or both configurations. 
       FIG. 2  is a cross section along line  2 - 2  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, about 25 to 55 about mils (about 635 to about 1400 microns) when the engine is cold to 0 to about 35 mils (about 889 microns) during engine operation depending on specific operations and previous rub events that may have occurred. 
     The new rotor coating is strong enough to abrade the bare super alloy vane tips by themselves thereby eliminating necessity of an abradable coating. 
       FIG. 2  and  FIG. 3  show 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. 
     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 . 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. 
     Bond coats  62  and  72  are thin, up to 10 mils, more specifically ranging from about 3 mils to about 7 mils (76 to 178 microns). Abrasive coatings  66  and  76  are much thicker than bond coats  62  and  72 , ranging from about 10 mils to about 19 mils (254 to 483 microns). 
     Bond coats  62  and  72  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 coats  62  and  64  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. Bond coat layers  62  and  72  are applied by plasma spraying. 
     Abrasive layer  66  and  76  may be a porous or filled metallic or ceramic material such as SM2042, SM2043, Metco 105NS or Durabrade 2192 available from Sulzer Metco. SM2042 is described in U.S. Pat. No. 5,434,210, which is incorporated by reference herein in its entirety. The selection of suitable abrasive layer material varies with application and is typically a compromise between erosion resistance, wear ratio with vane or blade tips and durability in the subject environment. One example choice may be Metco 105NS aluminum oxide coating with a mechanically roughened surface in an application where low erosion rate of the coating is desired. 
     Examples of Yttria stabilized zirconia layers  66  and  76  and metal bond coats  62  and  72  are described in commonly owned U.S. Pat. No. 5,879,753 and included herein in its entirety by reference. Coatings  66  and  76  in this patent consist essentially of zirconia containing 11-14 wt. % yttria. Coatings  66  and  76  are applied by plasma spraying, followed by laser engraving to form pyramids  66   a  and  76   a  on the surface facing the airfoil, as seen in  FIG. 3  and  FIG. 5 . 
     Other ceramic coatings may be used, provided that the ceramic has a coating having a hardness of 7 or higher on the Mohs scale of mineral hardness. These may be selected from quartz, zirconia such as those discussed above, corundum and diamond. 
       FIG. 6  and  FIG. 7  are enlarged photographs of pyramids  66   a  and  76   a  of  FIG. 3  and  FIG. 5 . Pyramids  66   a  and  76   a  are formed by application of a laser engraving on the surface that will engage the airfoil. The pyramids  66   a  and  76   a  were formed using a IPG 20W Q-switched fiber laser with a Nutfield XLR8-10-YAG 2-axis Scan Head with an f-theta 100 mm lens providing a max spot size of 16 μm. The f-theta 100 mm lens alloed for a working distance of 3.85″ in length. 
       FIG. 6  represents a finely spaced grit pattern with grit spacing of 0.005″, a texture height of 0.0015″, grit side slope of 45 degrees with respect to the surface before laser treatment, and the grits are misaligned in the circumferential direction.  FIG. 7  represents a coarsely spaced grit pattern with grit spacing of 0.010″, texture height of 0.0015″, grid side slope of 45 degrees and grits misaligned in the circumferential direction. 
     The pattern in  FIG. 6  was made at a power of 10 W, a speed of 900 mm/s, and a frequency of 40 kHz. Line length was 70 mm, line width was 0.05 mm, hatch distance was 0.004 mm and line distance was 0.381 mm. The pattern in  FIG. 7  was made at a power of 10 W, at a speed of 500 m/s, and a frequency of 40 kHz. Line length was 63.5 mm, line width of 0.126 mm, hatch distance of 0.004 mm and line distance of 0.381 mm. 
     The laser beam melts and removes parts of the ceramic coatings  66  and  76  at an angle with respect to the plane of the rotor or shroud so that the metallic airfoil encounters a sharp edge and is abraded. Other laser systems and dimensions are within the scope of this invention. In order to produce an effective grit surface using the laser treatment of the ceramic surface, the laser engraved surface with the top of the surface has less than about 5% of the surface area of the base of the ceramic layer. The degree of misalignment of the rows of pyramids can range from 0° to about 90°. The pyramids  66   a  and  76   a  of  FIGS. 6 and 7  are at a misalignment of 45° with respect to the circumferential direction of rotation about centerline C L . As seen in  FIGS. 6 and 7 , pyramids  66   a  and  76   a  form rows that are placed there by the laser action and the rows can be selectively aligned or misaligned as desired. 
     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.

Technology Category: f