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.

This article describes the visual, fractographic, and metallographic evidence typically encountered when analyzing stress rupture of turbine airfoils. Stress-rupture fractures are generally heavily oxidized, tend to be rough in texture, and are primarily intergranular and/or interdendritic in appearance compared to smoother, transgranular fatigue type fractures. Often, gross plastic yielding is visible on a macroscopic scale. Commonly observed microstructural characteristics include creep voiding along grain boundaries and/or interdendritic regions. Internal voids can also nucleate at carbides and other microconstituents, especially in single crystal castings that do not possess grain boundaries.

Aluminide-coated and uncoated IN-713 turbine blades were returned for evaluation after service in a marine environment because of severe corrosion. Based on service time, failure of these blades by corrosive deterioration was considered to be premature. Analysis (visual inspection, 2.7x micrographic examination on sections etched with ferric chloride and hydrochloric acid in methanol) supported the conclusions that the blades failed by hot-corrosion attack. Variation in rate of attack on coated blades was attributed to variation in integrity of the aluminide coating, which had been applied in 1966, when these coatings were relatively new. It is evident that maintaining the integrity of a protective coating could significantly increase the life of a nickel-base alloy blade operating in a hot and corrosive environment.

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.

Cracking in gas turbine blades was found to initiate from a mechanism of low-cycle fatigue (LCF). LCF is induced during thermal loading cycles in gas turbines. However, metallography of two cracked blades revealed a change in microstructure at as-cast surfaces for depths up to 0.41 mm (0.016 in.). Evaluation by SEM confirmed the difference in structure was associated with a lack of formation of coarse gamma prime structure in the matrix. Microhardness and miniature tensile test results indicated lower strength consistent with the absence of the coarse gamma prime constituent. The blade vendor found that the lot of hot isostatically pressed (HIP) blade castings had been exposed to an improper atmosphere during the HIP process, resulting in the weakened structure. Because subsequent failures were found in blades that did not come from the suspect HIP lot, the scope of the problem was considered generic, and the conclusion was that the primary failure mechanism was LCF. Material imperfections were a secondary deficiency that had the effect of causing the blades from the bad HIP lot to crack first.

Airfoil-shape impingement cooling tubes were fabricated of 0.25 mm (0.010 in.) thick Hastelloy X sheet stock, then pulse-laser-beam butt welded to cast Hastelloy X base plugs. Each weldment was then inserted through the base of a hollow cast turbine blade for a jet engine. The weldments were finally secured to the bases of the turbine blades by a brazing operation. One of the laser beam attachment welds broke after a 28-h engine test run. Exposure of the fracture surface for study under the electron microscope revealed the joint had broken in stress rupture. Failure was caused by tensile overload from stress concentration at the root of the laser beam weld, which was caused by the sharp notch created by the lack of full weld penetration. Radiographic inspection of all cooling-tube weldments was made mandatory, with rejection stipulated for joints containing subsurface weld-root notches. In addition, all turbine blades containing cooling-tube weldments were reprocessed by back-brazing. Back brazed turbine blades were reinstalled in the engine and withstood the full 150-h model test run without incident.