An 18-MW gas turbine exploded unexpectedly after three hours of normal operation. The catastrophic failure caused extensive damage to the rotor, casing, and nearly all turbo-compressor components. Based on their initial review, investigators believed that the failure originated at the interface between two shaft sections held together by 24 marriage bolts. Visual and SEM examination of several bolts revealed extensive deterioration of the coating layer and the presence of deep corrosion pits. It was also learned that the bolts were nearing the end of their operating life, suggesting that the effects of fatigue-assisted corrosion had advanced to the point where one of the bolts fractured and broke free. The inertial unbalance produced excessive vibration, subjecting the remaining bolts to overload failure.

Several surgical tool failures were analyzed to understand why they occur and how to prevent them. The study included drills, catheters, and needles subjected to the rigors of biomedical applications such as corrosive environments, high stresses, sterilization, and improper cleaning procedures. Given the extreme conditions to which surgical tools can be exposed, and the potential for misuse, failures are inevitable and systematic methods for analyzing them are necessary to keep them in check.

Two slitting saw blades were delivered for the purpose of determining the cause of damage. One had cracked while the other one came from a prior sheet delivery, that had less tendency to crack formation according to the manufacturer. The blades were supposed to have been stamped out of a sheet made from a 55 kp/sq mm strength steel. The saw blades were used for separating steel profiles at high rotational speeds. The cracks in question were located at the base of the teeth, i.e. at the point of highest operating stress. Metallographic examination showed that all cracks were non-decarburized and were free of chromium deposits. Therefore they could not have existed before heat treatment and chrome plating. It was concluded that the damage was due neither to poor quality of the sheet nor to defective stamping or heat treatment, but had occurred later either during surface treatment or during operation.

U-shaped leaf springs, intended to serve as spacers between oil tank floats and the inner walls of the containers, broke while being fitted, or after a short time in use, in the bend of the U. The springs were made of tempered strip steel of type C 88 with 0.84 % C, bent at room temperature, and electroplated with cadmium for protection against corrosion. Each fracture showed seven or eight kidney-shaped cracks. At the origins of these cracks on the concave inner surface of the springs, crater-like depressions and beads of melted and resolidified material were found. Fracture of the springs was caused by stress cracks as a consequence of local hardening. The hardening caused by melting and resolidification, and therefore the cracks in the springs, was the result of a faulty procedure during cadmium electroplating.

After only a short time in service, oil-fired orchard heaters made of galvanized low-carbon steel pipe, 0.5 mm (0.020 in.) in thickness, became sensitive to impact, particularly during handling and storage. Most failures occurred in an area of the heater shell that normally reached the highest temperature in service. A 400x etched micrograph showed a brittle and somewhat porous metallic layer about 0.025 mm (0.001 in.) thick on both surfaces of the sheet. Next to this was an apparently single-phase region nearly 0.05 mm (0.002 in.) in thickness. The examination supported the conclusion that prolonged heating of the galvanized steel heater shells caused the zinc-rich surface to become alloyed with iron and reduce the number of layers. Also, heating caused zinc to diffuse along grain boundaries toward the center of the sheet. Zinc in the grain boundaries reacted with iron to form the brittle intergranular phase, resulting in failure by brittle fracture at low impact loads during handling and storage. Recommendation included manufacture of the pipe with aluminized instead of galvanized steel sheet for the combustion chamber.

A failed 25 x 32 mm (1 x 1 in.) cadmium-plated 1040 carbon steel countersunk head type nose gear door securing bolt with a common screwdriver slot was examined. Fracture originated at a thread root and propagated across the cross section. The topography of the fracture was excessively rough and more granular than would be expected from pure mechanical fatigue. This indicated an allied corrosion mechanism. Cracks other than the one leading to failure were observed. Metallographic examination of the bolt cross section showed many cracks typical of stress-corrosion damage. It was concluded that the bolt failed by a combination of SCC and fatigue. It was recommended that aerospace-quality fasteners meeting NAS 7104, NAS 7204, or NAS 7504 be used to replace the currently used fasteners.

A failed H-11 tool steel pylon attachment stud was found during a routine walk-around inspection. The stud exhibited gross localized corrosion pitting at several different areas on its surface. Light general rust was also evident. Severe pitting occurred near the fracture location. The fracture face contained evidence of intergranular SCC as well as ductile dimples. The protective coating was found to be an inorganic water-base aluminide coating having a coating thickness of 7.5 to 13 micron (0.3 to 0.5 mil). The coating was of a nonuniform mottled nature. It was concluded that the failure of the pylon attachment stud was caused by general corrosion followed by SCC. The stud was not adequately protected against corrosion by the coating. It was recommended that the coating be applied to a thickness of 38 to 75 micron (1.5 to 3 mil) to provide long-time corrosion resistance. The coating must be either burnished or cured at 540 deg C (1000 deg F) to provide cathodic protection to the steel. Other coatings, such as cadmium or aluminum, were also recommended if a thinner coating is needed.

Forged aluminum alloy 2014-T6 hinge brackets in naval aircraft rudder and aileron linkages were found cracked in service. The cracks were in the hinge lugs, adjacent to a bushing made of cadmium-plated 4130 steel. Investigation (visual inspection and 250X micrographs) supported the conclusion that the failure of the hinge brackets occurred by SCC. The corrosion was caused by exposure to a marine environment in the absence of paint in stressed areas due to chipping. The stress resulted from the interference fit of the bushing in the lug hole. Recommendations included inspecting all hinge brackets in service for cracks and for proper maintenance of paint. Also suggested was replacing the aluminum alloy 2015-T6 with alloy 7075-T6, and surface treatment for the 7075-T6 brackets was recommended using sulfuric acid anodizing and dichromate sealing. Finally, it was also recommended that the interference fit of the bushing in the lug hole be discontinued.

Two clevis-head self-retaining bolts used in the throttle-control linkage of a naval aircraft failed on the aircraft assembly line. Specifications required the bolts to be heat treated to a hardness of 39 to 45 HRC, followed by cleaning, cadmium electroplating, and baking to minimize hydrogen embrittlement. The bolts broke at the junction of the head and shank. The nuts were, theoretically, installed fingertight. The failure was attributed to hydrogen embrittlement that had not been satisfactorily alleviated by subsequent baking. The presence of burrs on the threads prevented assembly to finger-tightness, and the consequent wrench torquing caused the actual fractures. The very small radius of the fillet between the bolt head and the shank undoubtedly accentuated the embrittling effect of the hydrogen. To prevent reoccurrence, the cleaning and cadmium-plating procedures were stipulated to be low-hydrogen in nature, and an adequate post plating baking treatment at 205 deg C (400 deg F), in conformity with ASTM B 242, was specified. A minimum radius for the head-to-shank fillet was specified at 0.25 mm (0.010 in.). All threads were required to be free of burrs. A 10-day sustained-load test was specified for a sample quantity of bolts from each lot.

Cartridge-pneumatic starter systems are used on military aircraft. In the cartridge mode used for alert starts, the starter turbine is driven by hot gases produced through the controlled burning of a solid propellant cartridge within a closed chamber (the breech chamber/cartridge chamber assembly). Premature failures of steel breech chambers have been prevalent enough to cause serious concern. The breech chamber is fabricated from a 4340 Ni-Cr-Mo steel forging heat treated to a hardness in the range HRC 40 to 45. The failures have taken several forms, including fracture and unzipping of the chamber dome, burn-through of the dome, and shearing of bayonet locking lugs. Factors identified as significant in the failures are the pressure developed in the chamber and internal corrosion of the chamber in an environment that can produce stress-corrosion cracking. The interior configuration of the chamber and the stress distribution also have a bearing upon the failure modes. Several failures are reviewed to illustrate the problems.

Ground maintenance personnel discovered hydraulic fluid leaking from two small cracks in a main landing gear cylinder made from AISI 4340 Cr-Mo-Ni alloy steel. Failure of the part had initiated on the ID of the cylinder. Numerous cracks were found under the chromium plate. A 6500x electron fractograph showed cracking was predominantly intergranular with hairline indications. Leaking had occurred only 43 h after overhaul of the part. Total service time on the part was 9488 h. It was concluded that cracking on the ID was caused by hydrogen embrittlement which occurred during or after overhaul. The specific source of hydrogen which produced failure was not ascertainable.

This report covers case histories of failures in fixed-wing light aeroplane and helicopter components. A crankshaft of AISI 4340 Ni-Cr-Mo alloy steel, heat treated and nitrided all over, failed in bending fatigue. The nitrided layer was ground too rapidly causing excessive heat generation which induced grinding cracks and grinding burn. Tensional stresses resulting from grinding developed in a thin surface layer. On another crankshaft, chromium plating introduced undesirable residual tensile stresses. Such plating is an unsatisfactory finish for crankshafts of aircraft engines. Aircraft engine manufacturers and aeronautical standards require magnetic particle inspection to detect grinding cracks after reconditioning. Renitriding after any grinding is needed also, regardless of the amount of undersize as it introduces beneficial residual compressive stresses.

A broken aircraft crankshaft and a severely damaged main brass bearing were examined to determine whether engine failure was initiated in the bearing or in the crankshaft. The steel crankshaft failure was a classical fatigue fracture. The bearing had been subjected to extremely high temperatures, as indicated by melting in the brass components and the extreme distortion in the rollers. Microscopic examination on the crankshaft material showed it to be a good quality steel. On the other hand, the chromium plate was thick, porous, and cracked in many places, including the point of the main fatigue crack. It was concluded that the over-all failure was initiated in the crankshaft, and the failure of the bearing resulted from that failure.

A large four-engine aircraft was on a cargo flight at night when a loud bang was heard, accompanied by a loss of power from both engines on the left side. After an emergency landing, it was discovered that the propellers from both left side engines were missing. The initial investigation determined that the four-bladed propeller from the left inboard engine had separated in flight, subsequently impacting the left outboard engine, causing its propeller to separate also. Three years later, the left inboard propeller hub was recovered. All four blades had separated through the shank area adjacent to the hub. Detailed SEM examination confirmed a fatigue mode of failure in this area with a primary single origin on the inside surface of the shank. The main fatigue origin site was coincident with one of the defects on the inner surface of the blade shank. The most probable source for creating the defects on the ID bore of the shank was the blade tip chrome plating process, which was carried out during the last overhaul prior to the failure.

Brittle intergranular fracture, typical of a hydrogen-induced delayed failure, caused the failure of an AISI 4340 Cr-Mo-Ni landing gear beam. Corrosion resulting from protective coating damage released nascent hydrogen, which diffused into the steel under the influence of sustained tensile stresses. A second factor was a cluster of non-metallic inclusions which had ‘tributary’ cracks starting from them. Also, eyebolts broke when used to lift a light aircraft (about 7000 lb.). The bolt failure was a brittle intergranular fracture, very likely due to a hydrogen-induced delayed failure mechanism. As for the factors involved, cadmium plating, acid pickling, and steelmaking processes introduce hydrogen on part surfaces. As a second contributing factor, both bolts were 10 Rc points higher in hardness than specified (25 Rc), lessening ductility and notch toughness. A third factor was inadequate procedure, which resulted in bending moments being applied to the bolt threads.

The outer tube, or stem, on a bicycle frame fractured after two years of use. Detailed investigation revealed that the lower stem bearing had been loose for some time and the bottom bearing cup contained many cracks. Metallographic examination of the chromium-plated cup confirmed the brittle nature of the cracks, located along prior austenite boundaries. The failure was attributed to hydrogen embrittlement due to improper manufacturing procedures following chromium plating. The cracking led to looseness in the bearing and consequent scoring, cracking, and overloading of the stem.

After less than 30 days in service, several cadmium-plated retaining rings, made of 4140 steel tubing and heat treated to 36 to 40 HRC, broke during operation that included holding components of a segmented fitting in place under a constant load. Photographic and 100x nital-etched micrographic examination showed a microstructure of tempered martensite with low inclusion content as well as a pit or burned spot on the outer area of the ring. The defect was approximately 0.18 mm (0.007 in.) deep and 0.5 mm (0.020 in.) in diam and had a hardness of 58 to 60 HRC. The base metal adjacent to the defect had a hardness of 36 to 40 HRC. Small cracks or fissures were also evident within the defect. Thus, the rings failed in brittle fracture as the result of an arc strike (or burn) on the surface of the ring. At the site of the arc strike, a small region of hard, brittle untempered martensite was formed as the result of an arc strike during the cadmium-plating operation. Fracture occurred readily when the ring was stressed. No recommendations were made.

Electroless nickel plating separation from copper alloy CDA175 retaining clips used on printed circuit boards was caused by a copper oxide layer that reduced adhesion of the nickel plating on the clips. Stresses that developed during module insertion caused flaking to occur at the oxidized copper surface. Electroless nickel plating separation from OFHC copper leads was caused by improper handling rather than a plating anomaly per se. Tin plating separation from copper underplating on a hybrid package lid occurred because of a four-week delay between the copper plating and tin plating steps. It was recommended that tin plating should follow the copper underplating within 24 h and a cleaning step of bright dipping after copper plating be performed.

Two types of chromium-plated hydraulic cylinders failed by cracking on their outer surfaces. In one case, the parts had a history of cracking in the nominally unstressed, as-fabricated condition. In another, cracks were detected after the cylinders were subjected to a pressure impulse test. Both part types were made of 15-5 PH (UNS S15500) precipitation hardening stainless steel. Hydrogen embrittlement cracking was the likely cause of failure for both part types. Cracking of the as-fabricated parts was ultimately prevented by changing the manufacturing procedure to allow for a reheat treatment. For parts that cracked after pressure testing, excessive dimensional changes precluded the inclusion of a reheat treatment as a manufacturing step, and further failure was averted by carefully employing proper machining practices, avoiding abusive machining.

A few Cr-Mo steel piston rods from different production batches were found identically cracked in the eye end near the radius after chrome plating and baking treatment. Two of them cracked in the plating stage itself instantly broke on slight tapping. Cracking initiated from the outer base surface of the forked eye end. The 40 mm diam forged piston rods were subjected to plating after heavy machining on the part without any stress-relieving treatment. Also, time lapses between plating and baking were varied from 3 to 11 h. The brittle cracking along forked eye-end radius portion was attributed to hydrogen embrittlement that occurred during chrome plating.