Simulation of Combustor Damage Mechanisms and Material System Performance via a Sub-Element Configured Specimen Test

Abstract

Modern gas turbine combustors are made of high temperature alloys, employ effusion cooling and are protected by a Thermal Barrier Coating (TBC). Standard material characterization tests such as creep, oxidation and low cycle fatigue are indicators of a material’s potential performance but they neither fully represent the combustor geometric/material system nor fully represent the thermal fatigue conditions a combustor is subjected to during engine operation. Combustor rig tests and/or engine cyclic endurance tests to determine the suitability of new material systems for combustors are time consuming and costly. Therefore, a simple test method for screening material systems under representative combustor conditions is needed. This experimental system was recently developed at Honeywell Aerospace to characterize various gas turbine combustor damage mechanisms and assess state-of-the-art and developmental materials. A configured specimen is fabricated using materials and processes similarly to actual combustor hardware, including sheet metal forming, welding, TBC coating, and effusion hole laser drilling. The configured specimen is cyclically exposed to hot spot thermal gradients typically experienced by fielded hardware using a jet-fueled burner and heated cooling air. Damage mechanisms simulated include bond coat oxidation, TBC spallation, thermal fatigue and distortion. A summary of these damage mechanisms and lessons learned from test development are presented. Results from recent combustor liner, bond coat, and top coat material modifications are also discussed. The effect of combustor liner material creep and thermal fatigue resistance, bond coat composition and processing, and TBC composition and structure on combustor durability is presented.