Gas turbine engines work as a power generating facility and are used in aviation industry to provide thrust by converting combustion products into kinetic energy [1-3]. Basic concerns regarding the improvements in modern gas turbine engines are higher efficiency and per‐ formance. Increase in power and efficiency of gas turbine engines can be achieved through increase in turbine inlet temperatures [1,4]. For this purpose, the materials used should have perfect mechanical strength and corrosion resistance and thus be able to work under aggressive environments and high temperatures [2]. The temperatures that turbine blades are exposed to can be close to the melting point of the superalloys. For this reason, internal cooling by cooling channels and insulation by thermal barrier coatings (TBCs) is used in order to lower the temperature of turbine blades and prevent the failure of superalloy substrates [1-4]. By utilizing TBCs in gas turbines, higher turbine inlet temperatures are allowed and as a result an increase in turbine efficiency is obtained [5]. TBCs are employed in a variety of areas such as power plants, advanced turbo engine combustion chambers, turbine blades, vanes and are often used under high thermal loads [6-11]. Various thermal shock tests are conducted by aerospace and land gas turbine manufacturers in order to develop TBCs and investigate the quality control characteristics. Despite that fact, a standardized method is still lacking. The reason lies behind the difficulty of finding a testing method that can simulate all the service and loading conditions. Present testing systems developed by the engine manufacturers for simulation of real thermal conditions in engines consist of; burner rig thermal shock testing units, jet engine thermal shock testing units and furnace cycle tests [16-20]. In this study, thermal cycle and thermal shock behavior of TBC systems under service conditions are examined, and a collection of testing methods used in evaluation of performance and endur‐