A primary constraint preventing the application of ceramic matrix composites in advanced gas turbine engines is a lack of adequate data and predictive modeling techniques to substantiate life limiting mechanisms such as creep, fatigue, and thermal shock induced rupture.
A two-dimensional transient heat transfer model was developed to calculate temperatures during the heating and cooling of hot pressed SCS-6 and SCS-9 SiC fiber reinforced ceramic composite specimens in the NASA Lewis Mach 0.3 atmospheric pressure burner test rig. The specimens modeled were thermally cycled under an impinging jet fuel flame in the temperature range 500 C to 1350 C under a constant applied tensile stress. An implicit finite difference procedure was employed that included the effect of forced convection, natural convection, and thermal radiation between the burner flame and the specimen.
The predicted temperature specimens are used to predict thermal stresses in the specimens.
"Heat Transfer During Burner Rig Thermal Fatigue of Ceramic Matrix Composites", T. Erturk, J. McKelliget, Ceram. Eng. Sci. Prod., 16,1995, pp. 95-104.
"Thermomechanical Fatigue of Crossply Si3/N4 Ceramic Composites", G. St. Hilaire, D. Eng. Thesis, University of Massachusetts at Lowell, Department of Mechanical Engineering, 1993.
© 1996, J. McKelliget