The primary purpose of this article is to describe general root causes of failure that are associated with wrought metals and metalworking. This includes a brief review of the discontinuities or imperfections that may be common sources of failure-inducing defects in the bulk working of wrought products. The article addresses the types of flaws or defects that can be introduced during the steel forging process itself, including defects originating in the ingot-casting process. Defects found in nonferrous forgings—titanium, aluminum, and copper and copper alloys—also are covered.

A forging die in a 250-ton press producing brass valves began to show signs of fatigue after a few thousand hits. By the time it reached 30,000 hits, the die was badly damaged and was submitted for analysis along with one of the last forgings produced. The investigation included visual and macroscopic inspection, metallographic and chemical analysis, SEM imaging, optical profilometry, mechanical property testing, and EDX analysis. The die was made of chromium hot-work tool steel and the forgings were made of CuZn39Pb3 heated to an initial working temperature 700 deg C. The entire surface of the die was covered with fatigue cracks and many fillets had been plastically deformed. Several other types of damage were also observed, including areas of oxidation, corrosion pits, voids, abrasive wear, die adhesion, and thermal fatigue. Fatigue cracking was the primary cause of failure with significant contributions from the other damage mechanisms.

A stainless steel screw securing an orthopedic implant fractured and was analyzed to determine the cause. Investigators used optical and scanning electron microscopy to examine the fracture surfaces and the microstructure of the austenitic stainless steel from which the screw was made. The results of the study indicated that the screw failed due to fatigue fracture stemming from surface cracks generated by stress concentration likely caused by grooves left by improper machining.

A deformed steel tube was received for failure analysis after buckling during a heat-treat operation. The tube was subjected to various metallurgical tests as well as nondestructive testing to confirm the presence of residual stresses. The microstructure of the tube was found to be homogenous and had no banded structure. However, x-ray diffraction analysis confirmed the presence of up to 6% retained austenite which likely caused the tube to buckle during the 910 °C heat treating procedure.

Several stainless steel coils cracked during a routine unwinding procedure, prompting an investigation to determine the cause. The analysis included optical and scanning electron microscopy, energy-dispersive x-ray spectrometry, and tensile testing. An examination of the fracture surfaces revealed a brittle intercrystalline mode of fracture with typical manifestations of clear grain facets. Branched and discrete stepwise microcracks were also found along with unusually high levels of residual hydrogen. Mechanical tests revealed a marked loss of tensile ductility in the defective steel with elongations barely approaching 8%, compared to 50% at the time of delivery weeks earlier. Based on the timing interval and the fact that failure occurred at operating stresses well below the yield point of the material, the failure is being attributed to hydrogen-induced damage. Potential sources of hydrogen are considered as are remedial measures for controlling hydrogen content in steels.

A pipe in the lateral wall of a boiler powering an aircraft carrier flat-top boat failed during a test at sea. The pipe was made from ASTM 192 steel, an adequate material for the application. Microstructural analysis along with equipment operating records provided valuable insight into what caused the pipe to rupture. Although the pipe had been replaced just 50 h before the accident, the analysis revealed incrustations and corrosion pits on the inner walls and oxidation on the outer walls. Microstructural changes were also observed, indicating that the steel was exposed to high temperatures. The combined effect of pitting, incrustations, and phase transformations caused the pipe to rupture.

A routine examination on a seat ejection system found that the catapult attachment swivel fabricated from 7075-T651 aluminum alloy plate contained cracks on opposite sides of the part. This swivel, or bath tub, does not experience extreme loads prior to activation of the catapult system. Some loads could be absorbed however, when the aircraft is subjected to G loads. Visual examination of the part revealed that cracks through the wall thickness initiated on the inner walls of the fixture. Scanning electron microscopy (SEM) and electron optical examination revealed that the cracking pattern initiated and progressed by an intergranular failure mechanism. It was concluded that failure of the catapult attachment swivel fixture occurred by SCC. It was recommended that the 7075 aluminum ejection seat fixture be supplied in the T-73 temper to minimize susceptibility to SCC.

A weld in a fuel-line tube broke after 159 h of engine testing. The 6.4-mm (0.25-in.) OD x 0.7-mm (0.028-in.) wall thickness tube and the end adapters were all of type 347 stainless steel. The butt joints between tube and end adapters were made by automated gas tungsten arc (orbital arc) welding. It was found that the tube had failed in the HAZ. Examination of a plastic replica of the fracture surface in a transmission electron microscope established that the crack origin was at the outer surface of the tube. The crack growth was by fatigue; closely spaced fatigue striations were found near the origin, and more widely spaced striations near the inner surface. The quality of the weld and the chemical composition of the tube both conformed to the specifications. However, the fuel-line assembly had vibrated excessively in service. The fuel-line fracture was caused by fatigue induced by severe vibration in service. Additional tube clamps were provided to damp the critical vibrational stresses. No further fuel-line fractures were encountered.

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.

Inspections and microstructural analysis revealed intergranular corrosion of 6061-T6 aluminum alloy aircraft fuel line beneath ferrules. The cause of the corrosion was traced to the fuel line marking process, which involved electrolytic labeling. Although subsequent rinsing of the fuel lines washed off most of the electrolyte, some was trapped between the 6061-T6 tubing and the ferrule. This condition made intergranular corrosion of the fuel lines inevitable. The attack caused grains to become dislodged, giving the appearance of pitting. Corrosion penetrated approximately 0.13 mm (0.005 in.) into the tubing. Experiments indicated that the corrosion products were inactive. It was recommended that another marking process be used that does not involve corrosive materials. The prevention of electrolyte from being trapped between the tubing and ferrules by using a MIL-S-8802 sealant was recommended.

On 16 July 1999, a Boeing 737-800 on final approach for landing sustained a major lightning strike. Damage to the fuselage structure primarily was in the form of melting or partial melting of widely-separated rivets and adjacent Alclad 2024-T3 fuselage skin. The damage was confined to a 0.25-in. (6.4-mm) radii around the affected rivets. The repair process involved removal of the locally-affected material and addition of a skin doubler to restore the aircraft structure to the originally designed condition. Damage features are described briefly.

Several elbow subassemblies comprising segments of oil-line assemblies that recycled aircraft-engine oil from pump to filter broke in service. The components of the subassemblies were made of aluminum alloy 6061-T6. Two subassemblies were returned to the laboratory to determine cause of failure. In one, the threaded boss had separated from the elbow at the weld. In the other, the failure was by fracture of the elbow near the flange. The separation of the threaded boss from the elbow was due to a poor welding procedure. Crack propagation was accelerated by fatigue caused by cyclic service stresses. The fracture of the second elbow near the flange was caused by overaging during repair welding of the boss weld. Satisfactory weld penetration was achieved by improved training of the welders plus more careful inspection. Repair welding was prohibited, to avoid recurrence of overaging from the welding heat. Additional support for the oil line was installed to reduce vibration and minimize fatigue of the elbow.

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.

Postflight inspection of a gas-turbine aircraft engine that had experienced compressor stall revealed that the engine air-intake bullet assembly had dislodged and was seated against the engine-inlet guide vanes at the 3 o'clock position. The bullet assembly consisted of an outer aerodynamic shell and an inner stiffener shell, both of 1.3 mm (0.050 in.) thick aluminum alloy 6061-T6, and four attachment clips of 1 mm (0.040 in.) thick alclad aluminum alloy 2024-T42. Each clip was joined to the outer shell by 12 spot welds and was also joined to the stiffener. Analysis (visual inspection, dye-penetrant inspection, and 10x/150x micrographs of sections etched with Keller's reagent) supports the conclusion that the outer shell of the bullet assembly separated from the stiffener because the four attachment clips fractured through the shell-to-clip spot welds. Fracture occurred by fatigue that initiated at the notch created by the intersection of the faying surfaces of the clip and shell with the spot weld nuggets. The 6061 aluminum alloy shell and stiffener were in the annealed (O) temper rather than T6, as specified. Recommendations included heat treating the shell and stiffener to the T6 temper after forming.

A series of resistance spot welds joining Z-shape and C-shape members of an aircraft drop-tank structure failed during ejection testing. The members were fabricated of alclad aluminum alloy 2024-T62. The back surface of the C-shape members showed severe electrode-indentation marks off to one side of the spot weld, suggesting improper electrode contact. Visual examination of the weld fractures showed that the weld nuggets varied considerably in size, some being very small and three exhibiting an HAZ but no weld. Of 28 welds, only nine had acceptable nugget diameters and fusion-zone widths. The weld deficiencies were traced to problems in forming and fit-up of the C-shape members and to difficulties in alignment and positioning of the weld tooling. The failure of the resistance spot welds was attributed to poor weld quality caused by unfavorable fit-up and lack of proper weld-tool positioning. The problem could be solved by better forming procedures to provide an accurate fit-up that would not interfere with electrode alignment.

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.

Two cracks were discovered in a deck plate of an aircraft during overhaul and repair after 659 h of service. The cracks were on opposite sides of the deck plate in the flange joggles. The plate had been formed from 7178-T6 aluminum alloy sheet. Analysis (visual inspection, 0.2x/2x/2.3x electron microscope fractographs, hardness testing, and electrical conductivity testing) supported the conclusions that the failure was caused by fatigue cracks originating on the inside curved surface of the flanges. The cracks had initiated in surface defects caused by either corrosion pitting or forming notches, acting in combination with lateral forces evidenced by the moderate distortion of the fastener holes. Recommendations included eliminating the surface defects by revised cleaning and/or forming procedures. Revised design and installation should also alleviate the lateral forces.

Resistance spot welds joining aluminum alloy 2024-T8511 stiffeners to the aluminum alloy 6061-T62 skin of an aircraft drop tank failed during slosh and vibration testing. Visual examination of the fracture surfaces showed that the failure was by tensile or bending overload. Measurements of the fractured spot welds established that all welds were below specification size. Review of the assembly procedures revealed that there had been poor fit-up between the stiffeners and the tank skin, which resulted in weak, undersize weld nuggets. The spot welds failed because of undersize nuggets that were the result of shunting caused by poor fit-up. The forming procedures were revised to achieve a precise fit between the stiffener and the tank wall. Also, an increase in welding current was suggested.

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.

Countersunk riveted joints in aluminum sheet are widely employed in the aircraft industry. The preparation of the sheet for the riveting process consists either of countersinking where the sheet is sufficiently thick or of dimpling. Metallographic assessment of dimple defects is described in specimens made of clad aluminum sheet of alloy type AlZnMgCu1.5. Addressed are a dimple with partially missing stamped surface (bell-mouth), a cylindrical prominence because the dimpling force was too great and the stamping cylinder force too low, and a dimple with flashes at the top surfaces of the sheet as a result of play between the stamping cylinder and the anvil head (ringed dimple). Frequently, overlapping of several defects occurs, especially with steel or titanium sheet, with the result that it is difficult to identify the defects.