Abstract:
A method for joining a first CMC part ( 30 ) having an outer joining portion ( 32 ), and a second CMC part ( 36 ) having an inner joining portion ( 38 ). The second CMC part ( 36 ) is heat-cured to a stage of shrinkage more complete than that of the first CMC part ( 30 ) prior to joining. The two CMC parts ( 30, 36 ) are joined in a mating interface that captures the inner joining portion ( 38 ) within the outer joining portion ( 32 ). The assembled parts ( 30, 36 ) are then fired together, resulting in differential shrinkage that compresses the outer joining portion ( 32 ) onto the inner joining portion ( 38 ), providing a tightly pre-stressed joint. Optionally, a refractory adhesive ( 42 ) may be used in the joint. Shrinkage of the outer joining portion ( 32 ) avoids shrinkage cracks in the adhesive ( 42 ).

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation-in-part of U.S. patent application Ser. No. 10/245,528, filed Sep. 17, 2002, now U.S. Pat. No. 7,093,359, incorporated by reference herein. 
    
    
     FIELD OF THE INVENTION 
     This invention relates generally to the field of ceramic materials, and more particularly to a method of joining ceramic elements. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engines are known to include a compressor section for supplying a flow of compressed combustion air, a combustor section for burning a fuel in the compressed combustion air, and a turbine section for extracting thermal energy from the combustion air and converting that energy into mechanical energy in the form of a shaft rotation. Many parts of the combustor section and turbine section are exposed directly to the hot combustion gasses, for example the combustor, the transition duct between the combustor and the turbine section, and the turbine stationary vanes, rotating blades and surrounding ring segments. 
     It is also known that increasing the firing temperature of the combustion gas may increase the power and efficiency of a combustion turbine. Modern high efficiency combustion turbines have firing temperatures in excess of 1,600 degrees C., which is well in excess of the safe operating temperature of the structural materials used in the hot gas flow path components. Special super alloy materials have been developed for use in such high temperature environments, and these materials have been used with specific cooling arrangements, including film cooling, backside cooling and insulation. 
     Ceramic and ceramic matrix composite (CMC) materials offer the potential for higher operating temperatures than do metal alloy materials, due to the inherent refractory nature of ceramic materials. This capability may be translated into a reduced cooling requirement that, in turn, may result in higher power, greater efficiency, and/or reduced emissions from the engine. 
     Prior art ceramic turbine airfoil members may be formed with an associated shroud or platform member. The platform defines a flow path between adjacent airfoil members for directing the hot combustion gasses past the airfoil members. The platform is exposed to the same high temperature gas environment as the airfoil member and thus may be formed of a ceramic material. The platform and the airfoil members may be formed as separate components that are unconnected and are allowed to have relative movement between them. However, such designs may not adequately transfer aerodynamic torque loads from the airfoil to the platform attachments. Alternatively, the platform and the airfoil may be formed as separate components that are mechanically joined together, as illustrated in U.S. Pat. No. 5,226,789. Such mechanical joints must be robust, and thus tend to be complicated and expensive. 
     Another alternative for joining the airfoil and the platform is to form the platform and the airfoil as a single integral part. Monolithic ceramic is readily moldable to a form, but it is limited to small shapes and is insufficiently strain-tolerant for robust designs. CMC materials incorporate ceramic fibers in a ceramic matrix for enhanced mechanical strength and ductility. However, conventional ceramic composite processing methods increase in complexity and cost in a complex three-dimensional component such as a turbine vane. U.S. Pat. No. 6,200,092 describes a turbine nozzle assembly having a vane forward segment formed of CMC material wherein the reinforcing fibers are specially oriented across the juncture of the airfoil and the platform members. Such special fiber placement in the airfoil-to-platform transition region presents a manufacturing challenge, especially with insulated CMC construction. Furthermore, for some CMC compositions, shrinkage during processing may result in residual stresses in complex shapes that are geometrically constrained. The airfoil-to-platform attachment area is one area where such stresses would arise. Additionally, load transfer between the airfoil and the platform results in interlaminar stresses in the fillet region where mechanical properties may be compromised. 
     In one solution to these problems, U.S. Pat. No. 6,648,597 discloses a method of manufacture for a vane component of a gas turbine, including: forming an airfoil member of a ceramic matrix composite material; forming a platform member of a ceramic matrix composite material; and forming an integral vane component by bonding respective joint surfaces of the airfoil member and the platform member. The method may further include: forming the airfoil member of a ceramic matrix composite material in a green body state; forming the platform member of a ceramic matrix composite material in a green body state; and urging the respective joint surfaces of the airfoil member and the platform member together at a firing temperature to form a sinter bond between them. The method may include densifying the sinter bond with a matrix infiltration process. The method may further include reinforcing the sinter bond with a fastener connected between the respective joint surfaces. Alternatively, the method may include bonding the respective joint surfaces of the airfoil member and the platform member with an adhesive. However, ceramic joints using refractory adhesives alone are weak and unreliable for carrying primary loads (mechanical, unrelenting loads). Furthermore, when such adhesives are applied to already-fired CMC parts in constrained geometries, the adhesives shrink and produce bond joint cracking. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in following description in view of the drawings that show: 
         FIG. 1  is a schematic sectional view of a joint in ceramic parts having a first mating geometry according to aspects of the invention; 
         FIG. 2  is a sectional view of a second mating geometry providing a bilateral shrinkage preload; 
         FIG. 3  is a sectional view of a third mating geometry with greater interlocking achieved by forming the outer joining part on the inner joining part; 
         FIG. 4  is a top view of a multi-part flange joined to a cylinder using clamping force provided by heat-curing shrinkage of a compression ring; 
         FIG. 5  is a sectional view taken along line  5 - 5  of  FIG. 4 ; 
         FIG. 6  is a perspective view of a flange with an outer joining portion segmented by slots 
         FIG. 7  is a sectional view of two parts interlocked by an intermediate insert and clamped by shrinkage of a compression ring; 
         FIG. 8  is a perspective view of a gas turbine vane joined to a platform member using compression rings; 
         FIG. 9  is a top view of the gas turbine vane and platform member of  FIG. 8 ; 
         FIG. 10  is an exploded view of the gas turbine vane and platform member  FIG. 8 ; 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  shows a schematic sectional view of a joint between a first ceramic part  30  and a second ceramic part  36 . The first ceramic part  30  has an outer joining portion  32  with an inner surface  34 . The second ceramic part  36  has an inner joining portion  38  with an outer surface  40  that mates with the inner surface  34  of the outer joining portion  32 . The second ceramic part  36  is heat-cured to a stage of shrinkage more complete than that of the first ceramic part  30  prior to joining. This means that both of these parts may be partially heat-cured to different stages, or the first ceramic part  30  may be left in a green body state and the second ceramic part  36  may be partially or fully cured. The term “green body state” includes stages of processing from a wet preform to a semi-fired state in which parts are rigid enough to be self-supporting. The outer joining portion  32  may be formed onto the inner joining portion  38 , such as by using the outer surface  40  of the inner joining portion  38  as a mold for lay-up of the first ceramic part  30  as a ceramic matrix composite (CMC) member. The assembly is then fully heat-cured. The first ceramic part  30  shrinks to a greater percentage than the second ceramic part  36 . This causes differential shrinkage  44 ,  46  of the outer joining portion  32  relative to the inner joining portion  38 , which tightens the outer joining portion  32  on the inner joining portion  38 , producing a pre-stressed joint. This method is especially useful for joining a first ceramic part  30  of CMC to a second ceramic part  36  of CMG, monolithic ceramic, and/or ceramic insulation. 
     A refractory adhesive  42  may be applied in the joining step. This fills any clearance between the outer joining portion  32  and the inner joining portion  38 . For example a high temperature ceramic adhesive such as from Cotronics® (Resbond 989 or 904), Aremco® (Ceramabond 503, 600, or 516), Sauerizon® (phosphate based adhesives), or Zircar® (ZR-COM) or variations on these basic adhesive types may be used. Shrinkage cracking in refractory adhesives is greatly reduced in the present method, because the outer joining portion  32  shrinks concurrently with the adhesive  42  and radially follows and compresses the adhesive. If an adhesive  42  or other filler is not used, the joining portions  32  and  38  may be formed and/or machined to minimize clearance between them. The outer joining portion  32  may be formed by CMC layering or continuous wrapping of ceramic fibers onto the inner joining portion  38  using the inner joining portion  38  as a form or mandrel for the outer joining portion  32  to eliminate all clearance there between. The fibers are impregnated with a ceramic matrix material, either before or after the wrapping step, to form a ceramic matrix composite material. 
     In  FIGS. 1-3  the outer joining portion  32  and inner joining portion  38  have mating geometries that retain the inner joining portion  38  within the outer joining portion  32  in at least two dimensions, X and Y relative to the drawing, after the joining step. In  FIGS. 1-3  the first ceramic part  30  can be provided as a single part. In this case, the inner joining portion  38  slides into the outer joining portion  32  along the Z dimension in the joining step. Alternately, the first ceramic part  30 , or at least the outer joining portion  32 , can be provided in two or more sections or slotted segments as later shown. These sections are first spread, then closed, around the inner joining portion  38 , and are then fixed by a compression ring  54  as next shown. With such segmented assembly, the mating geometries of  FIGS. 1-3  retain the inner portion  38  within the outer portion in three dimensions after the joining step. 
       FIGS. 4 and 5  illustrate a ceramic cylinder  50  joined with a ceramic segmented flange  52  held in place by a compression ring  54 , such as a CMC material. The compression ring  54  can be separately formed and applied in the green body state or it can be filament-wound around the outer joining portion  32  of the flange  52  in a wet state. In the latter case, drying shrinkage of the winding adds to the firing shrinkage to achieve an even greater amount of preload. The segmented flange  52  can be formed of multiple pieces as in  FIGS. 4 and 5 , or in one piece with slots  56  in the outer joining portion  32  as in  FIG. 6 , to allow compliance. 
       FIG. 7  shows an embodiment using a segmented flange  52  with an interlocking geometry including an intermediate locking ring  57 . This locking ring  57  can be applied as wet ceramic paste to both the inner surface  34  of the outer joining portion  32  and the outer surface  40  of the inner joining portion  38  prior to joining. It fills annular depressions in both of these surfaces, and merges into a single ring of material during joining and firing. Shrinkage in the locking ring  57  material during firing is accommodated by shrinkage in the compression ring  54 , reducing internal cracking in the locking ring  57  material. The locking ring  57  eliminates any possibility of separation or relative slippage between the parts in the direction of the longitudinal axis of cylinder  50 . If the flange is a multi-part type, the locking ring  57  can be pre-formed as a multi-part ring for assembly around the inner joining portion  38  during the joining step. 
     The ceramics utilized for the CMC materials in the joined parts may be oxide or non-oxide materials, for example alumina, mullite, silicon carbide, etc. The CMC compression ring  54  windings can be made with fibers other than those in the cylinder  50  or flange  52  pieces. Fibers with higher shrinkage, such as Nextel® 610 fibers (alumina) can be used in the compression ring versus Nextel® 720 fibers (alumina/mullite) typically used for the other parts. The higher shrinkage of the compression ring fibers imparts an even greater amount of prestress to the joint. Furthermore, the alumina fibers exhibit a greater coefficient of thermal expansion than do the alumina/mullite fibers, thereby imparting an additional amount of prestress as a result of the differential thermal expansion between the two types of fibers. The amount of prestress created by the drying shrinkage, the firing shrinkage and the thermal expansion of the mating parts can be selected to achieve a desired degree of prestress for both cold and hot conditions for any particular application. In one test embodiment, it was found that the combination of stresses imposed by drying and firing shrinkage and by differential thermal expansion was actually too high and resulted in an overstress condition in a CMC material containing Nextel® 610 fibers that were wet wrapped around a fully fired CMC member containing Nextel® 720 fibers. In order to reduce the level of prestress in such an embodiment, the Nextel® 610 fibers could be wet wrapped around a green body CMC member containing Nextel® 720 fibers, for example. 
       FIGS. 8-10  show how the present joining method might be applied to an airfoil-to-platform assembly. A gas turbine vane airfoil  70  may be formed comprising a core  72  and a skin  74 . For example, the core  72  may be a monolithic ceramic, and the skin  74  may be CMC and/or an insulating ceramic layer. One or both ends of the vane airfoil  70  may be attached to a respective platform member  76  by the present method. Outer joining portions are provided on the platform member in the form of tabs  78  extending from an opening  82  in the platform, as shown, to receive inner joining portions on an end of the vane airfoil  70 . In this embodiment, inner joining portions are provided on the ends of the vane airfoil  70  in the form of bosses  80 , as shown, formed in this example by removing parts of the ends of the vane airfoil  70 . The bosses  80  on the vane airfoil  70  are placed into the outer joining tabs  78  of the platform member  76 , and are clamped by differential shrinkage of compression rings  54  as previously described. One or more boss may be used and each boss may be adjoined to one or two opposed tabs in various embodiments. A filler material  58  may be inserted in some or all of the gaps between the compression rings  54  and the clamped parts  78  and  80 . Alternately, the inner joining bosses  80  and outer joining tabs  78  may be formed or machined with cylindrical side surfaces to match the inner surface of the compression rings  54 , thus avoiding some or all of the filler  58 . A refractory adhesive  42  may be used as previously described. 
     The vane airfoil  70  and/or the platform member  76  may be produced from a plurality of individual parts that are bonded or joined together in any variety of ways, or they may be formed as single-piece parts. The platform member  76  may be formed in two or more parts that are joined together during the joining step for the present method, with or without additional joining 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.