US9068464
(Ab)
A method for joining（接合）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（接合）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（接合部）. Shrinkage of the outer joining portion (32) avoids shrinkage cracks in the adhesive (42).
"BACKGROUND OF THE INVENTION
Gas turbine engines are known to include（～を有する～が公知である、知られている）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（を超える）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（冷却構成、手段）, 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（比較の対象：than metal alloy meterials offer the potential for operating temperatures? 厳密には"than the operating temperatures (またはthose) offered by metal alloy materials"？), due to the inherent refractory nature of ceramic materials. This capability may be translated into（この性能を利用して～する）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（強固）, 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（困難を伴う）, 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."
