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.

During a routine shear-pin check, the end lug on the barrel of the forward canopy actuator on a naval aircraft was found to have fractured. The lug was forged from aluminum alloy 2014-T6. Investigation (visual inspection, 2x views, and 140X micrographs etched with Keller's reagent) supported the conclusion that the cause of failure was SCC resulting from exposure to a marine environment. The fracture occurred in normal operation at a point where damage from pitting and intergranular corrosion acted as a stress raiser, not because of overload. The pitting and intergranular attack on the lug were evidence that the surface protection of the part had been inadequate as manufactured or had been damaged in service and not properly repaired in routine maintenance. Recommendations included anodizing the lug and barrel in sulfuric acid and giving them a dichromate sealing treatment, followed by application of a coat of paint primer. During routine maintenance checks, a careful examination was suggested to look for damage to the protective coating, and any necessary repairs should be made by cleaning, priming, and painting. Severely corroded parts should be removed from service.

A forged aluminum alloy 2014-T6 catapult-hook attachment fitting (anodized by the chromic acid process to protect it from corrosion) from a naval aircraft broke in service. Spectrographic analysis, visual examination, microscopic examination, and tensile analysis showed minute cracks on the inside surface of a bearing hole, and small areas of pitting corrosion were visible on the exterior surface of the fitting. The analysis also revealed a small number of rosettes, suggestive of eutectic melting, in an otherwise normal structure. These examinations and analyses support the conclusion that the presence of chromic acid stain on the fracture surface proved that the forging had cracked before anodizing. This suggest that the crack initiated during straightening, either after machining or after heat treatment. The structure and composition of the alloy appear to have been acceptable. Ductility was acceptable so rosettes found in the microstructure are believed to have been nondamaging. Had they contributed to the failure, the ductility would have been very low. The recommendations included inspection for cracks and revising the manufacturing process to include a fluorescent liquid-penetrant inspection before anodizing, because chromic acid destroys the penetrant. This inspection would reduce the possibility of cracked parts being used in service.

Four nose wheels fabricated from 2014-T6 aluminum alloy and cold worked at the flange were examined. Visual examination showed that the failure started in the tube well area on the wheel with serial number 31. The failure initiated in the flange fillet on wheels with serial numbers 67, 217, and 250. Scanning electron microscopy (SEM) examination of the fractures showed that failure initiated by SCC or a corrosion pit on all failures examined. The failures then progressed by fatigue. Dye penetrant testing revealed no additional flaws on the wheels that had failed in the flange area. There was, however, one flaw area in the flange of the wheel that failed in the tube well. This flaw resembled a corrosion pit. It was concluded that failure of nose wheels 67, 217, and 250 was caused by cracking due to SCC or pitting. The failures progressed by fatigue. Because failure occurred in the same general area on all three wheels, these locations are suspect as being underdesigned. It was recommended that the nose wheel be redesigned and additional service data be accumulated to understand the contributing factors that resulted in wheel failure.

The flange on an outboard main-wheel half (aluminum alloy 2014-T6 forging) on a commercial aircraft fractured during takeoff. The failure was discovered later during a routine enroute check. The flange section that broke away was recovered at the airfield from which the plane took off and was thus available for examination. Failure occurred after 37 landings (about 298 roll km, or 185 roll miles). Examination of the fracture surfaces revealed that a forging defect was present in the wall of the wheel half. The anodized coating showed distinct twin-parallel and end-grain patterns between which the fracture occurred. The periphery of the defect was the site of several small fatigue cracks that eventually progressed through the remaining wall. Rapid fatigue then progressed circumferentially. Metallographic examination using Keller's reagent showed that the microstructure was normal for aluminum alloy 2014-T6 and the hardness surpassed the minimum hardness required for aluminum alloy 2014-T6. An abrupt change in the direction of grain flow across the fracture plane indicated that the wall had buckled during forging. This evidence supported the conclusion that the wheel half failed in the flange by fatigue as the result of a rather large subsurface forging defect. No recommendations were made.

During a routine inspection, cracks were discovered in several aluminum alloy (similar to either 2014 or 2017) coupling nuts on the fuel lines of a missile. The fuel lines had been exposed to a marine atmosphere for six months while the missile stood on an outdoor test stand near the seacoast. A complete check was then made, both visually and with the aid of a low-power magnifying glass, of all coupling nuts of this type on the missile. Investigation (visual inspection, spectrographic and chemical analysis, and metallographic examination) supported the conclusion that the cracking of the aluminum alloy coupling nuts was caused by stress corrosion. Contributing factors included use of a material that is susceptible to this type of failure, sustained tensile stressing in the presence of a marine (chloride-bearing) atmosphere, and an elongated grain structure transverse to the direction of stress. The elongated grain structure transverse to the direction of stress was a consequence of following the generally used procedure of machining this type of nut from bar stock. Recommendations included changing the materials specification for new coupling nuts for this application to permit use of only aluminum alloys 6061-T6 and T651 and 2024-T6, T62, and T851.

A commercial aircraft wheel half, machined from an aluminum alloy 2014 forging that had been heat treated to the T6 temper, was removed from service because a crack was discovered in the area of the grease-dam radius during a routine inspection. Neither the total number of landings nor the roll mileage was reported, but about 300 days had elapsed between the date of manufacture and the date the wheel was removed from service. The analysis (visual inspection, macrographs, micrographs, electron microprobe) supported the conclusions that the wheel half failed by fatigue. The fatigue crack originated at a material imperfection and progressed in more than one plane because changes in the direction of wheel rotation altered the direction of the applied stresses. Recommendations included rewriting the inspection specifications to require sound forgings.

Two outboard main-wheel halves (aluminum alloy 2014-T6 forged) from a commercial aircraft were removed from service because of failure. One wheel half was in service for 54 days and had made 130 landings (about 1046 roll km, or 650 roll mi) when crack indications were discovered during eddy-current testing. The flange on the second wheel half failed after only 31 landings, when about 46 cm (18 in.) of the flange broke off as the aircraft was taxiing. Stains on the fracture surfaces were used to determine when cracking was initiated. The analysis (visual inspection, liquid penetrant inspection, and micrographs with deep etching in aqueous 20% sodium hydroxide) supported the conclusion that failure on both wheel halves was by fatigue caused by a forging defect resulting from abnormal transverse grain flow. The crack in the first wheel half occurred during service, and the surfaces became oxidized. Because the fracture surface of the second wheel half had chromic acid stains, it was obvious that the forging defect was open to the surface during anodizing. No recommendations were made except to notify the manufacturer.

Proper stress analysis during component design is imperative for accurate life and performance prediction. The total stress on a part is comprised of the applied design stress and any residual stress that may exist due to forming or machining operations. Stress-corrosion cracking may be defined as the spontaneous failure of a metal resulting from the combined effects of a corrosive environment and the effective component of tensile stress acting on the structure. However, because of the orientation dependence in aluminum, it is the residual stress occurring in the most susceptible direction that must be considered of primary importance in material selection for design configuration. A Navy UH-1N helicopter main rotor blade grip manufactured from a 2014-T6 aluminum alloy forging failed because of a design flaw that left a high residual tensile stress along the short transverse plane; this in turn provided the necessary condition for stress corrosion to initiate. A complete failure investigation to ascertain the exact cause of the failure was conducted utilizing stereomicroscopic examination, scanning electron microscopy, metallographic inspection and interpretation, energy-dispersive chemical analysis, physical and mechanical evaluation. Stereomicroscopic examination of the opened crack fracture surface revealed one large fan-shaped region that had propagated radially through the thickness of the material from two distinct origin areas on the internal diam of the grip. Higher magnification inspection near the origin area revealed a flat, wood-like appearance. Scanning electron microscopy divulged the presence of substantial mud cracking and intergranular separation on the fracture surface. Metallographic examination revealed intergranular cracking and substantial leaf separation along the elongated grains parallel to the fracture surface. Chemical composition and hardness requirements were found to be as specified. The blade grip failed due to a stress corrosion crack which initiated on the inner diam and propagated in the short transverse direction through the thickness of the component. The high residual tensile stress in the part resulting from the forging and exposed after machining of the inner diam, combined with the presence of moisture, provided the necessary conditions to facilitate crack initiation and propagation.

An aircraft engine in which an in-flight fire had occurred was dismantled and examined. A bracket assembly fabricated from 2024 aluminum, one of several failed components, was of prime interest because of apparent heat damage. Scanning electron microscopy was used to compare laboratory-induced fractures made at room and elevated temperatures with the bracket failure. The service failure exhibited grain separation and loss of delineation of the grain boundaries due to melting. SEM revealed deep voids between grains and tendrils that connected grains, which resulted from surface tension during melting. Microscopic examination of polished, etched section through the fractured surface verified intergranular separation and breakdown of grain facets. The absence of any reduction of thickness on the bracket assembly at the point of fracture, along with evidence of intense heat at this point, indicated that little stress had been applied to the part. Comparisons of the service failure and laboratory-induced failures in conjunction with macroscopic and metallographic observations showed that the bracket assembly failed because an intense, localized flame had melted the material.

A helicopter tail rotor blade spar failed in fatigue, allowing the blade to separate during flight. The 2014-T652 aluminum alloy blade had a hollow spar shank filled with lead wool ballast and a thermoset polymeric seal. A corrosion pit was present at the origin of the fatigue zone and numerous trails of corrosion pits were located on the spar cavity's inner surfaces. The corrosion pitting resulted from the failure of the thermoset seal in the spar shank cavity. The seal failure allowed moisture to enter into the cavity. The moisture then served as an electrolyte for galvanic corrosion between the lead wool ballast and the aluminum spar inner surface. The pitting initiated fatigue cracking which led to the spar failure.

Two complete aircraft undercarriage-leg 2014 aluminum alloy forgings and a number of sectional ends that exhibited cracks during nondestructive testing were examined to determine the extent of damage and the type of cracking. Cracks were primarily confined to the diaphragm and adjoining wall between the steel sleeve and the steel diaphragm washer. Metallographic analysis and accelerated corrosion tests showed that the cracks had originated as stress-corrosion failures.

A piece of wheel flange separated from the main landing gear wheel of a C130 aircraft as it taxied on a runway. The wheel was a 2014-T61 aluminum alloy forging and had been in service nearly 20 years. Fractographic evidence indicated that the initial crack growth was caused by high-cycle fatigue. The crack grew to approximately 8 in. in length before final catastrophic fracture. Fatigue analyses accurately predicted the cyclic life demonstrated by the failed wheel since its last inspection, assuming an initial crack length of 13 to 25 mm (0.5 to 1.0 in.). It was recommended that the inspection interval be reduced to one-third of its original duration for the current level of inspection reliability, or that inspection procedures be improved in order that cracks substantially smaller than 13 mm (0.5 in.) can be reliably detected.

Despite extensive aircraft landing gear design analyses and tests performed by designers and manufacturers, and the large number of trouble-free landings, aircraft users have experienced problems with and failures of landing gear components. Different data banks and over 200 failure analysis reports were surveyed to provide an overview of structural landing gear component failures as experienced by the Canadian Forces over the last 20 years on more than 20 aircraft types, and to assess trends in failure mechanisms and causes. Case histories were selected to illustrate typical problems, troublesome failure mechanisms, the role of high strength aluminum alloys and steels, and situations where fracture mechanics analyses provided insight into the failures. The two main failure mechanisms were: fatigue occurring mainly in steel components, and corrosion related problems with aluminum alloys. Very few overload failures were noted. A number of causes were identified: design deficiencies and manufacturing defects leading mainly to fatigue failures, and poor materials selection and improper maintenance as the principal causes of corrosion-related failures. The survey showed that a proper understanding of the failure mechanisms and causes, by thorough failure analysis, provides valuable feedback information to designers, operators and maintenance personnel for appropriate corrective actions to be taken.