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.

Examination of several fighter aircraft main landing gear legs revealed unusual cracking in the hard chromium plating that covered the sliding section of the inner strut. The cracking was associated with cracks in the 35 NCD 16 steel beneath the plating. A detailed investigation revealed that the cracking was caused by the combination of incorrect grinding procedure, the presence of hydrogen, and fatigue. The grinding damage generated tensile stresses in the steel, which caused intergranular cracking during the plating cycle. The intergranular cracks were initiation sites for fatigue crack growth during service. It was recommended that the damaged undercarriage struts be withdrawn from service pending further analysis and development of a repair technique.

One main undercarriage axle made of high strength alloy steel was subjected to simulated fatigue test for 6000 h of service. After only 300 h it broke in two along the sharp radius. The fracture revealed a coarse, irregular, and brittle surface before final fracture by thick angular shear lip zone. The presence of micropores in the cleavage facets as well as at the grain boundaries and hairline type crack indications under SEM examination were all suggestive of hydrogen embrittlement. On the basis of investigation results and observations, it was concluded that the transverse breakage of the axle had occurred intergranularly in a brittle manner, possibly, initiated by a shallow zone of fatigue along the sharp radius acting as stress riser.

The truck beam of the left main landing gear (MGL) of a Boeing 707 airplane collapsed on the ground just after the aircraft was unloaded and refueled. The investigation revealed that failure was caused by the propagation of an intergranular crack originating from the bottom of the pit. The crack reached the critical size and caused failure by stress-corrosion cracking (SCC) under static loading conditions in service. The failed beam was protected by a well adhering paint system. However, the presence of adequate amounts of corrosion preventive compound films (CPC) on the surfaces of the failed beam could not be conclusively established because of the long term service exposure and presence of lubricants.

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.

Service failures have occurred in a number of aircraft parts made of quenched and tempered steel heat treated to ultimate tensile strengths of 260,000 to 280,000 psi. Some of these failures have been attributed to “delayed cracking” as a result of hydrogen embrittlement or to stress-corrosion. Because of the serious nature of the failures and because the mechanism of the fracture initiation is not well understood, unusually complete laboratory investigations have been conducted. Three of these investigations are reviewed to illustrate the methods used in studying failures in aircraft parts. The results of the laboratory studies indicate that unusual care is necessary in the processing and fabrication of ultra-high-strength steel and in the design and maintenance of the structures in which it is used.

A flat spring for the main landing gear of a light aircraft failed after safe execution of a hard landing. The spring material was identified by chemical analysis to be 6150 steel. The fracture was found to have occurred near the end of the spring that was inserted through a support member about 25 mm thick and attached to the fuselage by a single bolt. Brinelling (plastic flow and indentation due to excessive localized contact pressure) was observed on the upper surface of the spring where the forward and rear edges of the spring contacted the support member. It was indicated by chevron marks that brittle fracture had started beneath the brinelled area at the forward edge of the upper surface of the spring. The origin of the brittle fracture was found to be a small fatigue crack that had been present for a considerable period of time before final fracture occurred. Fracture of the landing-gear spring was concluded to have been caused by a fatigue crack that resulted from excessive brinelling at the support point. Regular visual examinations to detect evidence of brinelling and wear at the support in aircraft with this configuration of landing-gear spring were recommended.

A connecting rod from a failed engine ruptured in fatigue without evidence of excessive stresses, detonation, overheating, or oil starvation. The origin of the fatigue failure was completely mutilated but decarburization was observed. Significant amounts of decarburization (0.010 to 0.015 in.) were found also in other forgings, such as exhaust rocker arms, main rotor drag brace clevises, bolts of carriage diagonal struts, and spring legs of main landing gears. The failure mode was low-stress, high-cycle fatigue involving tension and bending loads. The main cause was a manufacturing deficiency. The usual way to eliminate decarburization is to machine off the soft skin or employ better quality control when making them. Many aircraft manufacturers employ forged parts with machined surfaces or with shot-peened as-forged surfaces without excessive decarburization.

Fretting and/or fretting corrosion fatigue have been observed on such parts as main rotor counterweight tie rods, fixed-pitch propeller blades, propeller blade clamps, pressure regulator lines, and landing gear support brackets. Microcracks started from severe corrosion pits in a failed control rotor spar tube assembly made of cadmium-plated AISI 4130 Cr-Mo alloy steel. Inadequate design was responsible for the failure. A lower tine of the main rotor blade cuff failed in fatigue. The rotor blade cuff was forged of 2014-T6 aluminum alloy. Initial stages of crack growth displayed features typical of low stress intensity fatigue of aluminum alloys. The fatigue resulted from abnormal fretting owing to inadequate torquing of the main retention bolts. Aircraft maintenance engineers and owners were advised to adhere to specifications when torquing this joint.

The US Army Research Laboratory performed a failure investigation on a broken main landing gear mount from an AH-64 Apache attack helicopter. A component had failed in flight, and initially prevented the helicopter from safely landing. In order to avoid a catastrophe, the pilot had to perform a low hover maneuver to the maintenance facility, where ground crews assembled concrete blocks at the appropriate height to allow the aircraft to safely touch down. The failed part was fabricated from maraging 300 grade steel (2,068 MPa [300 ksi] ultimate tensile strength), and was subjected to visual inspection/light optical microscopy, metallography, electron microscopy, energy dispersive spectroscopy, chemical analysis, and mechanical testing. It was observed that the vacuum cadmium coating adjacent to the fracture plane had worn off and corroded in service, thus allowing pitting corrosion to occur. The failure was hydrogen-assisted and was attributed to stress corrosion cracking (SCC) and/or corrosion fatigue (CF). Contributing to the failure was the fact that the material grain size was approximately double the required size, most likely caused from higher than nominal temperatures during thermal treatment. These large grains offered less resistance to fatigue and SCC. In addition, evidence of titanium-carbo-nitrides was detected at the grain boundaries of this material that was prohibited according to the governing specification. This phase is formed at higher thermal treatment temperatures (consistent with the large grains) and tends to embrittle the alloy. It is possible that this phase may have contributed to the intergranular attack. Recommendations were offered with respect to the use of a dry film lubricant over the cadmium coated region, and the possibility of choosing an alternative material with a lower notch sensitivity. In addition, the temperature at which this alloy is treated must be monitored to prevent coarse grain growth. As a result of this investigation and in an effort to eliminate future failures, ARL assisted in developing a cadmium brush plating procedure, and qualified two Army maintenance facilities for field repair of these components.

The right landing gear on a twin-turboprop transport aircraft collapsed during landing. Preliminary examination indicated that the failure occurred at a steel-to-aluminum (7014) pinned drag-strut connection due to fracture of the lower set of drag-strut attachment lugs at the lower end of the oleo cylinder housing. Two lug fractures that were determined to be the primary fractures were analyzed. Results of various examinations indicated that stress-corrosion cracking associated with the origins of the principal fractures in the connection was the cause of failure. It was recommended that the design be modified to avoid dissimilar metal combinations of high corrosion potential.

In a spring leg of a main landing gear, large brittle fracture zones indicated a predominately cleavage pattern with some ductile dimples, and a tiny fatigue segment disclosed fine striations. Factors influencing failure were surface decarburization, notch sensitivity of the modified SAE 6150 spring steel, Canada's cold weather which may have had an embrittling effect on the steel, and cumulative fatigue damage from severe landing loads during service life. Replacement with heavier-duty spring legs will probably not eliminate this type of failure, but their use has reduced the number of failures substantially. Precautionary measures recommended to preclude accidents include removal of decarburization, proper operation of main landing gears, and adequate magnetic particle inspection of the legs at the beginning and end of the ski season to detect any fatigue cracks that might develop in attachment holes.

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.

A crack was detected in one arm of the right-hand horizontal brace of the nose landing gear shock strut from a large military aircraft. The shock strut was manufactured from a 7049 aluminum alloy forging in the shape of a delta. A laboratory investigation was conducted to determine the cause of failure. It was concluded that the arm failed because of the presence of an initial defect that led to the initiation of fatigue cracking. The fatigue cracking grew in service until the part failed by overload. The initial defect was probably caused during manufacture. Fleet-wide inspection of the struts was recommended.

The failure of an ATAR engine accessory angle drive gear assembly caused an engine flame-out in a Mirage III aircraft of the Royal Australian Air Force (RAAF) during a landing. Stripping of the engine revealed that the bevel gear locating splines (16 NCD 13) had failed. Visual and low-power microscope examination of the spline of the shaft showed evidence of fretting wear debris; similar wear was observed on the splines of the mating bevel gear. It was concluded that the splines had failed by severe fretting wear. Fretting damage was also observed on the shaft face adjacent to the splines and on the bevel gear abutment shoulder. Additional tests included a metrological inspection of the shaft, bevel gear and support ring; metallographic examination of a section from the shaft; chemical analysis of the shaft material (16 NCD 13); and hardness testing of a sample of the yoke material. The wear had been caused by incorrect machining of the shaft splines, which prevented the bevel gear nut from locating correctly against the gear.

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 failures of two aircraft components, one from a landing gear and the other from an ejector rack mechanism, were investigated. Both were made from PH 13-8 Mo (UNS S13800) precipitation-hardening stainless steel which had been heat treated to the H1000 and H950 tempers respectively and then chromium plated. The parts were characterized metallographically and mechanically and were found to be compliant. Detailed fractographic examination revealed that the first stage of both failures was similar: subsurface initiation of numerous cracks with a wide range of orientations and cleavage like features. The cracking was followed by fatigue in one case and catastrophic failure in the other. Hydrogen embrittlement was identified as the most likely mechanism of failure.

This article focuses on failure analyses of aircraft components from a metallurgical and materials engineering standpoint, which considers the interdependence of processing, structure, properties, and performance of materials. It discusses methodologies for conducting aircraft investigations and inspections and emphasizes cases where metallurgical or materials contributions were causal to an accident event. The article highlights how the failure of a component or system can affect the associated systems and the overall aircraft. The case studies in this article provide examples of aircraft component and system-level failures that resulted from various factors, including operational stresses, environmental effects, improper maintenance/inspection/repair, construction and installation issues, manufacturing issues, and inadequate design.