New aircraft wing panels extruded from 7075-T6 aluminum exhibited an unusual pattern of circular black interrupted lines, which could not be removed by scouring or light sanding. The panels, subsequent to profiling and machining, were required to be penetrated inspected, shot peened, H2SO4 anodized, and coated with MIL-C-27725 integral fuel tank coating on the rib side. Scanning electron microscopy and microprobe analysis (both conventional energy-dispersive and Auger analyzers) showed that the anodic coating was applied over an improperly cleaned and contaminated surface. The expanding corrosion product had cracked and, in some places, had flaked away the anodized coating. The corrodent had penetrated the base aluminum in the form of subsurface intergranular attack to a depth of 0.035 mm (0.0014 in.). It was recommended that a vapor degreaser be used during cleaning prior to anodizing. A hot inhibited alkaline cleaner was also recommended during cleaning prior to anodizing. The panels should be dichromate sealed after anodizing. The use of deionized water was also recommended during the dichromate sealing operation. In addition, the use of an epoxy primer prior to shipment of the panels was endorsed. Most importantly, surveillance of the anodizing process itself was emphasized.

A crack was found in an aircraft main wing spar flange fabricated from 7079-T6 aluminum alloy during a routine nondestructive x-ray inspection after the craft had logged 300 h. Scanning electron microscopy (SEM) revealed an intergranular fracture pattern indicative of stress-corrosion cracking (SCC) and fatigue striations near the crack origin. Visual examination of the crack edge revealed that the installation of the fasteners produced a fit up stress. Further inspection of the opened fracture showed that the crack had been present for some time because a heavy buildup of corrosion products was seen on the fractured surface. Metallographic examination of the flange in the area of fracture initiation showed the presence of end grain exposure, which would promote SCC. Electron optical examination of the fracture clearly showed the flange was cracking by a mixed mode of stress corrosion and fatigue. The cracking was accelerated because of an inadvertent fit up stress during installation. The age of the crack could not be established. However, a reevaluation of prior x-ray inspections in this area would result in some close estimate of the age of the crack. End grain exposure further promoted SCC.

During a routine inspection on an aircraft assembly line, an airframe attachment bolt was found to be broken. The bolt was one of 12 that attach the lower outboard longeron to the wing carry-through structure. Failure occurred on the right-hand forward bolt in this longeron splice attachment. The bolt was fabricated from PH13-8Mo stainless steel heat treated to have an ultimate tensile strength of 1517 to 1655 MPa (220 to 240 ksi). A water-soluble coolant was used in drilling the bolt hole where this fastener was inserted. Investigation (visual inspection, 265 SEM images, hardness testing, auger emission spectroscopy and secondary imaging spectroscopy, tensile testing, and chemical analysis) supported the conclusion that failure of the attachment bolt was caused by stress corrosion. The source of the corrosive media was the water-soluble coolant used in boring the bolt holes. Recommendations included inspecting for corrosion all the bolts that were installed using the water-soluble coolant at the spliced joint areas, rinsing all machined bolt holes with a noncorrosive agent, and installing new PH13-8Mo stainless steel bolts with a polysulfide wet sealant.

Cracks were found on the wing leading edge of a test aircraft made from AZ31B magnesium alloy. Crack lengths were approximately 230 mm (9 in.) long on the left side and approximately 130 mm (5 in.) long on the right side. The cracks ran parallel to the leading edge. The 230-mm (9-in.) crack was received for examination. Visual examination of the submitted panel revealed two cracks. One crack ran through six adjacent fastener holes. Sections of the beveled edges of the holes were missing and corrosion was evident. Visual examination of the fastener holes after separation of the crack showed that the fracture faces were corroded. Optical examination of either side of the middle group of fastener holes showed that the area of suspected crack initiation had suffered excessive corrosion. Examination of the holes on the end of the crack showed fracture characteristics typical of fatigue and/or corrosion fatigue. It was concluded that crack propagation of the fracture in the wing panel occurred by a combination of corrosion and high-cycle fatigue in the end fastener holes. It was recommended that future panels be manufactured of 2024 aluminum.

Two titanium alloy wing attachment bolts from a commercial jetliner failed during the course of a routine service operation. Failure of the bolts occurred during the re-torque process as the wing was being reattached. Metallurgical failure analysis indicated that the fracture mechanism was ductile overload and that the mechanical properties of the bolts were consistent with exemplar bolts that had been supplied. After eliminating other sources of excessive load application, the most probable cause of failure was ascribed to variances between the frictional characteristics of the bolt at the time of re-torque and at the time of initial torque application several years earlier.

Following the crash of a Mirage III-0 aircraft (apparently caused by engine failure), a small crack was detected in a bolt hole in the wing main spar (AU4SG aluminum alloy). Because this area was considered to be critical to aircraft safety and similar cracking was found in other spars in service, the Royal Australian Air Force requested that the crack growth rate during service be determined. The loading history of the aircraft was made available in the form of flight by-flight records of the counts from the vertical accelerometer sensors fitted to the airframe and a series of “overstress” events recorded during the life of the aircraft. The bolt hole was examined by eddy current testing, visual examination, high-powered light microscope, and scanning electron microscope. Simulation tests were also conducted. The use of simulation specimens permitted actual crack growth rate data to be determined for the configuration of interest.

Cracks were discovered in the cast 17-4 PH stainless steel outboard leading edge flap support of an aircraft wing during overhaul inspection. Failure analysis focused on an apparently intergranular area of fracture surface. It was determined that the original mode of crack growth was cleavage, probably caused by cast-in hydrogen. The intergranular appearance resulted from heat treatment of the already cracked part, which caused the formation of grain-boundary “growth figures” on the exposed crack surfaces. It was recommended that the castings be more closely inspected for defects before further processing and that foundry practices be reviewed to correct deficiencies leading to excessive hydrogen absorption during melting and casting.