Patent Document

FIELD OF THE INVENTION 
     The present invention relates to abradable surfaces for high temperature applications, and more particularly relates to such surfaces on ceramic matrix composite (CMC) ring segments for combustion turbines. 
     BACKGROUND OF THE INVENTION 
     Some components of combustion turbines operate at high temperatures, and thus may require thermal barrier coatings (TBCs). Conventional TBCs typically comprise a thin layer of zirconia or other ceramic material. In some applications, the coatings must be erosion resistant and also abradable. Turbine ring seal segments must withstand erosion and must also have tight tolerances on a radially inward sealing surface opposed the tips of rotating turbine blades. To minimize these tolerances, the sealing surface of ring segments may be made abradable in order to reduce damage to the turbine blades upon occasional brushing contact of blade tips with the sealing surface. 
     Improvements in gas turbine efficiency rely on breakthroughs in several key technologies as well as enhancements to a broad range of current technologies. One of the key issues is a need to tightly control rotating blade tip clearance. This requires that turbine ring segments are able to absorb mechanical rubbing by rotating blade tips. 
     For modern conventionally cooled and closed loop steam cooled turbine ring segments, a thick thermal barrier coating of about 0.1 inch on the ring segment surface is required for rubbing purposes. The latest advanced gas turbine has a hot spot gas temperature of over 1,500 degrees C. at the first stage ring segment. Under such heat, a TBC surface temperature of over 1,300 degrees C. is expected. Thus, a conventional abradable TBC is no longer applicable because conventional TBCs are typically limited to a maximum surface temperature of about 1,150 degrees C. 
     Friable graded insulation (FGI) materials are disclosed in U.S. Pat. No. 6,641,907 commonly owned by the present assignee. The effectiveness of FGI as an abradable refractory coating is based upon control of macroscopic porosity in the FGI to deliver acceptable abradability. Such a coating may consist of hollow ceramic spheres in a matrix of alumina or aluminum phosphate. To bond an FGI layer to a metal ring segment, an FGI-filled metallic honeycomb structure has been proposed in U.S. Pat. No. 6,846,574 commonly owned by the present assignee. In this technique a high temperature metal alloy honeycomb is brazed to the metallic substrate. The honeycomb, once oxidized prior to FGI application, serves as a mechanical anchor and compliant bond surface for an FGI filler, and provides increased surface area for bonding. 
     Further advances in high temperature abradable surfaces for gas turbine ring segment surfaces are desired. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in the following description in view of the drawings that show: 
         FIG. 1  is a sectional and perspective view of a prior art CMC wall for a gas turbine ring segment with a thermal barrier coating. 
         FIG. 2  is a sectional and perspective view of a CMC wall for a gas turbine ring segment with depressions providing an abradable sealing surface according to the invention. 
         FIG. 3  shows CMC lamina having respective edges profiled with maxima and minima, and staggered in a stacked construction to form a CMC wall providing a sealing surface with depressions according to the invention. 
         FIG. 4  shows a variation of the lamina edge shapes of  FIG. 3 . 
         FIG. 5  shows CMC lamina in a first series having respective edges shaped with maxima and minima, a second series having respective edges that are generally level with the maxima of the first series, and the first and second series of the lamina alternating with each other in a stacked wall construction. 
         FIG. 6  shows an array of unconnected depressions, each with a generally circular opening geometry. 
         FIG. 7  shows an array of unconnected depressions, each with a generally rectangular opening geometry formed by lamina as in  FIG. 3 . 
         FIG. 8  shows an array of unconnected depressions, each with a generally hexagonal opening geometry. 
         FIG. 9  is a sectional and perspective view of a CMC wall for a gas turbine ring segment formed by stacked laminate construction with depressions providing an abradable sealing surface. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     A gas turbine component, especially a ceramic matrix composite (CMC) ring segment, is described herein with an abradable surface exposed to a hot gas flow. In contrast to prior art, no thermal barrier coating is applied to the exposed surface. Instead, the CMC itself is used as its own thermal barrier, but is modified to allow for abradability. The current invention provides an array of depressions directly in the CMC surface to increase its abradability, allowing occasional brushing contact with turbine blade tips with reduced wear on the blade tips. This technology is especially applicable to CMC ring segment walls formed by laminate construction, in which CMC layers are oriented edgewise in a stacked configuration. 
       FIG. 1  illustrates a CMC wall structure  22  of a prior art ring segment  20 P that has a thermal barrier coating  24  such as FGI to provide an abradable gas flow sealing surface  26 .  FIG. 2  illustrates a CMC wall structure  32  of a ring segment  30  that has a sealing surface  34  with no coating, but with an array of depressions  36  according to aspects of the invention to increase the abradability of the surface  34 . The depressions  36  are unconnected to each other in order to prevent bypass of the working gas around the blade tips via the depressions. They can be formed by removal of material from the CMC surface  34  after constructing and curing the wall  32 , or they can be formed by laminate edge profiling, as next described. Material removal processes may include one or more known methods, such as milling, drilling, water jet cutting, laser cutting, electron beam cutting, and ultrasonic machining. 
       FIG. 3  illustrates a CMC wall structure  48  formed by a stack of CMC layers (or lamellae)  40 - 43  with edge profiling  50 ,  52  that results in a surface  44  with unconnected depressions  46 . Techniques for manufacturing such a stacked lamellate assembly are known in the art, such as discussed in commonly-assigned United States Patent Application Publications US 2006/0121265 and US 2006/0120874, both incorporated by reference herein. Each layer  40 - 43  has a respective edge that is profiled with alternating maxima  50  and minima  52  that may be formed onto the edge prior to joining of the lamellae together. The maxima  50  and minima  52  are staggered in alternating layers  40 - 43  so that the adjoining maxima  50  of one or several adjacent layers are substantially aligned with the adjoining minima  52  of one or several adjacent layers to form a plurality of unconnected depressions  46  in the surface  44 . In the embodiment of  FIG. 3 , the maxima  50  and minima  52  both define a generally rectangular shape, with the relative absolute and relative depth and length dimensions of the rectangular shapes being selectable by the designer to optimize performance in any specific application. Typically, the dimensions of the depressions  46  may be 1.5-2.5 mm deep and up to 4 mm long (i.e. along the longitudinal axis of the lamella). Typically the length of the exposed maxima surface segment  50  may be 5-7 mm. 
       FIG. 4  illustrates a variation of  FIG. 3  in a CMC wall structure  48 ′ formed by a stack of CMC layers  40 ′- 43 ′ with edge profiling  50 ′,  52 ′ that results in a surface  44 ′ with unconnected depressions  46 ′. Each layer  40 ′- 43 ′ has a respective edge that is profiled with alternating maxima  50 ′ and minima  52  that define a generally V-shape. The dimensions of the exposed surface segment  50 ′ and the depth of the depression  46 ′ may be similar to those described for the embodiment of  FIG. 3 . 
       FIG. 5  illustrates a CMC wall structure  48 ″ formed by a stack of CMC layers  40 ″- 43 ″ in which a first series  40 ″ and  42 ″ of the CMC layers has maxima  50 ″ and minima  52 ″, a second series  41 ″ and  43 ″ of the layers has respective edges  53  that generally match the level of the maxima  50 ″ of the first series, and the first and second series of the layers  40 ″- 43 ″ alternates in the stack. In this embodiment the transition between the maxima  50 ″ and minima  52 ″ define a relatively smooth curved shape. The dimensions of the exposed surface segment  50 ″ and the depth of the depression  46 ″ may be similar to those described for the embodiment of  FIG. 3 . Other edge profiles and arrangements are possible. For example profiles similar to those of the first series of CMC layers  40 ″ and  42 ″ of  FIG. 5  could be used in a staggered configuration as in  FIGS. 3 and 4 , and vice versa. 
       FIG. 6  illustrates an array of unconnected depressions  36  with circular openings in a surface  34 , as may be formed by ball-end milling or other machining processes. The depressions  36  may have a spherical shape, or they may have a cylindrical shape proximate the surface  34  with a spherical bottom, or they may have a cylindrical shape throughout. Embodiments wherein depressions have a cross-sectional area that decreases with depth are effective to present an increasing wear surface area as the sealing surface is worn by abrasion, thereby facilitating the wear-in of the surface. 
       FIG. 7  illustrates an array of unconnected depressions  46  with rectangular openings in a surface  44  formed by a stacked laminate construction as in  FIG. 3 . 
       FIG. 8  illustrates an array of unconnected depressions  54  with hexagonal openings in a surface  34 ′, as may be formed by laser, water jet, or electron beam machining techniques. 
       FIG. 9  illustrates a turbine ring segment  30 ′ with a CMC wall  32 ′ formed by bonding and curing of stacked CMC lamellae  56 . A gas sealing surface  34 ″ on the wall  32 ′ is subsequently machined with an array of depressions  36  according to the invention as in  FIGS. 2 and 6  or in other shapes such as illustrated in  FIGS. 3-5 ,  7  and  8 . 
     Behavior of CMC exposed to high temperatures shows reduction in strength over long periods; however such a reduction in strength should not be limiting for the present invention because strength is not the material property of primary concern for a wear surface. Since a CMC surface  34 ,  44  in this invention is directly exposed to the hot working gas, it will be exposed to temperatures over 1200° C. This will reduce its strength but will also increase its hardness. The increase in hardness will beneficially reduce erosion of the surface. The surface may be allowed to age during operation of the gas turbine engine, or it may be pre-aged prior to being placed into operation. A thin, hard ceramic coating, for example alumina, may be applied to the CMC edges as temporary erosion protection until CMC hardening occurs. 
     The present invention eliminates the need for an abradable thermal barrier coating such as FGI, thus eliminating the associated bond joint and avoiding any concern about differential elasticity between the two materials. Accordingly, the invention is expected to provide improved component reliability and durability and reduced manufacturing expense compared to prior art coating methods. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.

Technology Category: 2