Coating system for clearance control in rotating machinery

Abstract

The invention relates to gas turbine engine seal systems having a rotating member with an abrasive tip surface disposed in rub relationship to a stationary seal member with an abradable surface. The abrasive tip surface is coated with a metallic alloy matrix having ceramic abrasive particles embedded in and projecting from the matrix. The abradable seal surface is coated with a ceramic coating. The rub relationship affords a tight operating clearance between the rotating member and the stationary seal member, thereby improving engine efficiency, reducing fuel consumption and minimizing overhaul downtime.

1 . A gas turbine engine seal system comprising a rotating member having an abrasive tip surface disposed in rub relationship to a stationary abradable seal surface, wherein said abrasive tip surface comprises an abrasive coating deposited onto at least a portion of the tip surface, wherein said abrasive coating comprises a metallic alloy matrix having ceramic abrasive particles at least partially embedded in said matrix and at least some of the ceramic abrasive particles projecting from said matrix, wherein said ceramic abrasive particles are selected from alumina polycrystal, alumina single crystal (sapphire), chromia-doped alumina single crystal (ruby), yttria-alumina garnet (YAG), titania-doped alumina polycrystal or single crystal (emerald), SiAlON, SiC, Si 3 N 4  or diamond; wherein said abradable seal surface comprises an abradable ceramic coating deposited onto at least a portion of the seal surface; and wherein said gas turbine engine seal system has a tip-to-seal wear ratio of at least about 1:10. 
     
     
         2 . A gas turbine engine seal system comprising a rotating member having an abrasive tip surface disposed in rub relationship to a stationary abradable seal surface, wherein said abrasive tip surface comprises an abrasive coating deposited onto at least a portion of the tip surface, wherein said abrasive coating comprises a metallic alloy matrix having ceramic abrasive particles at least partially embedded in said matrix and at least some of the ceramic abrasive particles projecting from said matrix; wherein said abradable seal surface comprises an abradable ceramic coating deposited onto at least a portion of the seal surface, wherein said ceramic coating is made from a ceramic powder comprising ceramic powder macroparticles, said ceramic powder macroparticles comprising a zirconia-based component and an (alumina+silica)-based component, wherein said ceramic powder macroparticles contain from about 10 to about 95 percent by weight of the zirconia-based component and about 5 to about 90 percent by weight of the (alumina+silica)-based component, and wherein the average particle size (diameter) of the ceramic powder macroparticles is from about 10 to about 150 microns; and wherein said gas turbine engine seal system has a tip-to-seal wear ratio of at least about 1:10. 
     
     
         3 . The gas turbine engine seal system of  claim 1  having a tip-to-seal wear ratio of at least about 1:20. 
     
     
         4 . The gas turbine engine seal system of  claim 1  wherein the metallic alloy matrix comprises MCrAlY where M is Ni, Co, Fe or combinations thereof. 
     
     
         5 . The gas turbine engine seal system of  claim 4  wherein the metallic alloy matrix is selected from NiCrAlY, NiCoCrAlY and CoNiCrAlY. 
     
     
         6 . The gas turbine engine seal system of  claim 1  wherein the ceramic abrasive particles comprise angular ceramic abrasive particles of nominal size 4 to 15 mils. 
     
     
         7 . The gas turbine engine seal system of  claim 1  wherein the ceramic abrasive particles have a hardness of from about 1000 Kg/mm 2  to about 7000 Kg/mm 2  and a fracture toughness of from about 1.5 Mpa*m 0.5  to about 8 Mpa*m 0.5 . 
     
     
         8 . The gas turbine engine seal system of  claim 1  wherein the ceramic abrasive particles are embedded in said matrix to a depth of about nominally half the size of the ceramic abrasive particles, with an upper portion of the ceramic abrasive particles projecting above said matrix. 
     
     
         9 . The gas turbine engine seal system of  claim 1  wherein a bondcoat is deposited between the tip surface and the abrasive coating or deposited between the seal surface and the ceramic coating or both. 
     
     
         10 . The gas turbine engine seal system of  claim 9  wherein the bondcoat comprises MCrAlY where M is Ni, Co, Fe or combinations thereof. 
     
     
         11 . The gas turbine engine seal system of  claim 9  wherein the bondcoat comprises: (i) an alloy containing chromium, aluminum, yttrium with a metal selected from the group consisting of nickel, cobalt and iron; or (ii) an alloy containing aluminum and nickel. 
     
     
         12 . The gas turbine engine seal system of  claim 9  wherein the bondcoat comprises a MCrAlY+X coating, where M is Ni, Co or Fe or any combination of the three elements, and X includes the addition of Pt, Ta, Hf, Re or other rare earth metals, or fine alumina dispersant particles, singularly or in combination. 
     
     
         13 . The gas turbine engine seal system of  claim 9  wherein the bondcoat has a surface roughness of at least about 150 microinches. 
     
     
         14 . The gas turbine engine seal system of  claim 1  which is heated in vacuum at a temperature sufficient to create a bond between the ceramic coating and the seal surface or between tip surface and the abrasive coating or both. 
     
     
         15 . The gas turbine engine seal system of  claim 1  wherein the abrasive coating is deposited by electroplating. 
     
     
         16 . The gas turbine engine seal system of  claim 1  wherein said abrasive coating thickness is from about 0.0025 to about 0.10 inches. 
     
     
         17 . The gas turbine engine seal system of  claim 1  wherein said abradable ceramic coating comprises a stabilized zirconia coating, a stabilized zirconia coating having a macrocracked microstructure, or a coating having a zirconia-based component and an (alumina+silica)-based component. 
     
     
         18 . The gas turbine engine seal system of  claim 1  wherein the abradable ceramic coating comprises (i) at least one metallic or metallic/ceramic inner layer deposited onto at least a portion of the seal surface, (ii) optionally at least one ceramic intermediate layer deposited onto the inner layer, and (iii) at least one abradable ceramic outer layer deposited onto the inner layer, or optionally the intermediate layer. 
     
     
         19 . The gas turbine engine seal system of  claim 18  wherein said intermediate layer has a plurality of macrocracks distributed throughout the intermediate layer. 
     
     
         20 . The gas turbine engine seal system of  claim 1  wherein said abradable ceramic coating is made from a ceramic powder that is blended with a fugitive material. 
     
     
         21 . The gas turbine engine seal system of  claim 20  wherein the fugitive material is selected from controlled size particles of a polyester or Lucite. 
     
     
         22 . The gas turbine engine seal system of  claim 20  wherein the ceramic powder and fugitive material are thermally sprayed to form a precursor coating that is heat treated to volatilize the fugitive material and to produce a coating having at least about 20% porosity. 
     
     
         23 . The gas turbine engine seal system of  claim 20  wherein the fugitive material is separately injected into a thermal spray device at a point of lower plume enthalpy and co-sprayed with the ceramic powder. 
     
     
         24 . The gas turbine engine seal system of  claim 1  wherein said abradable ceramic coating is made from a ceramic powder that is blended with a solid lubricant. 
     
     
         25 . The gas turbine engine seal system of  claim 24  wherein the solid lubricant is selected from hexagonal boron nitride. 
     
     
         26 . The gas turbine engine seal system of  claim 24  wherein the solid lubricant is separately injected into a thermal spray device at a point of lower plume enthalpy and co-sprayed with the ceramic powder. 
     
     
         27 . The gas turbine engine seal system of  claim 1  wherein the abradable ceramic coating has a porosity that increases from an inner surface of the ceramic coating to an outer surface of the ceramic coating. 
     
     
         28 . The gas turbine engine seal system of  claim 1  wherein the abradable ceramic coating has a density of from about 45 to about 90 percent of theoretical. 
     
     
         29 . The gas turbine engine seal system of  claim 1  wherein the abradable ceramic coating has a thickness of from about 0.02 to about 0.10 inches. 
     
     
         30 . The gas turbine engine seal system of  claim 1  wherein the abradable ceramic coating is deposited by plasma spray, detonation gun, high velocity oxy-fuel (HVOF), or high velocity air-fuel (HVAF). 
     
     
         31 . The gas turbine engine seal system of  claim 1  wherein infeed per strike of the abrasive tip surface to the stationary abradable seal surface generates wear debris particles having an average particle size (diameter) of from about 1 micron or less to about 150 microns. 
     
     
         32 . The gas turbine engine seal system of  claim 1  wherein the rotating member is a turbine blade. 
     
     
         33 . The gas turbine engine seal system of  claim 1  wherein the rotating member is a turbine rotor knife edge disposed on a turbine rotor and the abradable seal surface is disposed on a turbine vane to form an inner air seal. 
     
     
         34 . The gas turbine engine seal system of  claim 1  wherein the rotating member is a compressor blade. 
     
     
         35 . The gas turbine engine seal system of  claim 1  wherein the rotating member is a compressor rotor knife edge disposed on a compressor rotor and the abradable seal surface is disposed on a compressor stator to form an inner air seal. 
     
     
         36 . The gas turbine engine seal system of  claim 2  wherein the average particle size (diameter) of the ceramic powder macroparticles is from about 10 to about 100 microns. 
     
     
         37 . The gas turbine engine seal system of  claim 2  wherein said ceramic powder comprises a blend of said ceramic powder particles that is spray dried and optionally sintered to produce composite ceramic powder macroparticles. 
     
     
         38 . The gas turbine engine seal system of  claim 37  wherein said blend of ceramic powder particles is controllably spray dried and sintered to produce a size and distribution of microporosity within said composite ceramic powder macroparticles. 
     
     
         39 . The gas turbine engine seal system of  claim 2  wherein the abradable ceramic coating comprises two or more sublayers in which (i) the amount of the zirconia-based component and (alumina+silica)-based component continuously change throughout the sublayers, or (ii) the amount of the zirconia-based component and (alumina+silica)-based component discretely change from one sublayer to another. 
     
     
         40 . The gas turbine engine seal system of  claim 2  wherein the abradable ceramic coating further comprises a plurality of vertical macrocracks extending at least half the coating thickness in length up to the full thickness of the coating and having from about 5 to about 200 vertical macrocracks per linear inch measured in a line parallel to the surface of the seal and in a plane perpendicular to the surface of the seal. 
     
     
         41 . The gas turbine engine seal system of  claim 40  wherein said abradable ceramic coating has at least about 40 vertical macrocracks per linear inch measured in a line parallel to the surface of the seal and in a plane perpendicular to the surface of the seal. 
     
     
         42 . The gas turbine engine seal system of  claim 40  wherein the width of the vertical macrocracks is less than 1 mil.