Patent ID: 9102015
Filing Date: 2015-08-11
Classification: B22F,B23K,B23P,C22C,C23C,F01D,F05B,F05D,Y10T

Abstract:
1. A method for fabricating a thermal barrier of a gas turbine component, comprising: determining at least one material property of an existing thermal barrier coating on the component, the material property selected from the group consisting of modulus of elasticity, thermal conductivity, and coefficient of thermal expansion, wherein a value of said at least one material property is determined as a function of operating temperature and operating duration of the existing thermal barrier coating; and disposing a first layer of a metal brazing material in a die case; disposing a second layer of a ceramic material on the metal brazing material; and co-sintering the first and second layers in the die case with spark plasma sintering to form a co-sintered ceramic/metal tile comprising a metal brazing layer, a ceramic thermal barrier layer, and a co-sintered interface there between; and controlling the co-sintering to make the ceramic/metal tile to match the value of said at least one material property within 10% over an expected range of operating temperatures and time.