Gas turbines and other types of combustion turbomachinery are susceptible to hot corrosion at elevated temperatures. Two such cases resulting in the failure of a gas turbine component were investigated to learn more about the hot corrosion process and the underlying failure mechanisms. Each component was analyzed using optical and scanning electron microscopy, energy dispersive spectroscopy, mechanical testing, and nondestructive techniques. The results of the investigation provide insights on the influence of temperature, composition, and microstructure and the contributing effects of high-temperature oxidation on the hot corrosion process. Preventative measures are also discussed.

A microstructural analysis has been made of a burner nozzle removed from service in a coal gasification plant. The nozzle was a casting of a Co-29wt%Cr-19wt%Fe alloy. Extensive hot corrosion had occurred on the surface. There was penetration along grain boundaries, and corrosion products in these regions were particularly rich in S, and also contained Al, Si, O, and Cl. The grain boundaries contained Cr-rich particles which were probably Cr23-C6 type carbides. In the matrix, corrosion occurred between the Widmanstatten plates. Particles were found between these plates, most of which were rich in Cr and O, and probably were Cr2-O3 oxides. Other matrix particles were found which were rich in Al, O, and S. The corrosion was related to these grain boundary and matrix particles, which either produced a Cr-depleted zone around them or were themselves attacked.

The circumstances surrounding the in-service failure of a cast Ni-base superalloy (Alloy 713LC) second stage turbine blade and a cast and coated Co-base superalloy (MAR-M302) first stage air-cooled vane in two turbine engines used for marine application are described. An overview of a systematic approach, analyzing the nature of degeneration and failure of the failed components, utilizing conventional metallurgical techniques, is presented. The topographical features of the turbine blade fracture surface revealed a fatigue-induced crack growth pattern, where crack initiation had taken place in the blade trailing edge. An estimate of the crack-growth rate for the stage II fatigue fracture region coupled with the metallographic results helped to identify the final mode of the turbine blade failure. A detailed metallographic and fractographic examination of the air-cooled vane revealed that coating erosion in conjunction with severe hot-corrosion was responsible for crack initiation in the leading edge area.

A turbine vane made of cast cobalt-base alloy AMS 5382 (Stellite 31; composition: Co-25.5Cr-10.5Ni-7.5W) was returned from service after an undetermined number of service hours because of crack indications on the airfoil sections. This alloy is cast by the precision investment method. Analysis (visual inspection, 100x/500x metallographic examination of sections etched with a mixture of ferric chloride, hydrochloric acid, and methanol, and bend tests) supported the conclusions that cracking of the airfoil sections was caused by thermal fatigue and was contributed to by low ductility due to age hardening, subsurface oxidation related to intragranular carbides, and high residual tensile macrostresses. No further conclusions could be drawn because of the lack of detailed service history, and no recommendations were made.