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.

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.

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.

Aluminum 7075 aircraft wing tanks failed in the 1950s. Investigation (visual inspection, biological analysis, and chemical analysis) supported the conclusion that MIC was the cause of the failures. Water condensed into the fuel tanks during flight led to microbial growth on the jet fuel. Pitting attack occurred under microbial deposits on the metal surface in the water phase or at the water-fuel interface. Previously, exposure of aluminum 7075 to cultures of various isolates showed that 27 bacterial isolates and 3 fungi could seriously corrode the aluminum alloy over several weeks. No recommendations were made.

A longeron assembly constructed of Alclad 2024, some parts being in the T3 condition, others in the T42 condition, failed at a rivet hole. Plastic deformation at the crack site was found, but no plastic deformation was found in similar failed components. It was concluded that the numerous hairline cracks in the Alclad layer adjacent to the main fracture were fatigue cracks. In another case, bonded samples of 2024-T3 sheet were fatigue tested at various stress levels. Failures could be separated into three groups: those that failed in the adhesive bond, those that failed in the base material, and those that exhibited a dual failure. The last category failed in the adhesive bond and also showed a type of pitting on one face of the base material. In a third case, a 2024-T4 extrusion section was found to exhibit blistering after chemical milling. The presence of interconnecting microcracks between adjacent discontinuities supported a hydrogen blistering diagnosis.

The torque-arm assembly (aluminum alloy 7075-T73) for an aircraft nose landing gear failed after 22,779 simulated flights. The part, made from an aluminum alloy 7075-T73 forging, had an expected life of 100,000 simulated flights. Initial study of the fracture surfaces indicated that the primary fracture initiated from multiple origins on both sides of a lubrication hole that extended from the outer surface to the bore of a lug in two cadmium-plated flanged bushings made of copper alloy C63000 (aluminum bronze) that were press-fitted into each bored hole in the lug. Sectioning and 2x metallographic analysis showed small fatigue-type cracks in the hole adjacent to the origin of primary fracture. Hardness and electrical conductivity were typical for aluminum alloy 7075. This evidence supported the conclusion that the arm failed in fatigue cracking that initiated on each side of the lubrication hole since no material defects were found at the failure origin. Recommendations included redesign of the lubrication hole, shot peeing of the faces of the lug for added resistance to fatigue failure, and changing of the forging material to aluminum alloy 7175-T736 for its higher mechanical properties.

This paper summarizes the results of a failure analysis investigation of a fractured main support bridge made of 7075 aluminum alloy from an army helicopter. The part, manufactured by “Contractor IT,” failed component fatigue testing while those of the original equipment manufacturer (OEM) passed. Metallurgical data collected during this investigation indicated that the difference in fatigue life between the components fabricated by IT and by OEM may be attributable to a difference in dimensions at the web where fatigue crack initiation occurred. The webs of the two OEM parts examined had cross-sectional thicknesses significantly larger than the web cross-sectional thicknesses of the IT components. Recommendations included changing the web reference dimension of 0.38 in. to include a tolerance range based upon a fracture mechanics model. Also, the shot peening process should be controlled especially at the critical areas of the web, to assure complete coverage and proper compressive residual stresses.

An aluminum alloy propeller blade that had been cold straightened to correct deformation incurred in service fractured soon after being returned to service. Visual examination revealed that crack initiation occurred at the top surface in an area containing numerous surface pits. Macroscopic appearance of the surface was of brittle fracture. X-ray stress analysis did not detect any residual stress in the top surface of the propeller blade adjacent to the fracture. However, a spanwise tensile stress of approximately 51 MPa (7.4 ksi) was indicated in the same surface of the unfailed mating blade at the location of the initial bend. Evidence found supports the conclusions that the residual stress probably originated with straightening, and the apparent absence of stress in the fractured blade was the result of relaxation through fracture. Because no prior crack damage could be attributed to the initial deformation or to straightening, rapid fracture may have been induced by residual stresses contributing to the normal spectrum of cyclic stresses. Recommendations included stress-relief annealing after cold straightening, refinishing of the surface, thus reducing fracturing of propeller blades that were cold straightened to correct deformation experienced in service.

The floors (fabricated from aluminum alloy 7178-T6 sheet, with portions of the sheet chemically milled to reduce thickness) of the fuel tanks in two aircraft failed almost identically after 1076 and 1323 h of service, respectively. Failure in both tanks occurred in the rear chemically milled section of the floor. An alkaline etch-type cleaner was used on the panels before chemical milling and before painting. Various tests and measurements indicated that the aluminum alloy used for the fuel-tank floors conformed to the specifications for 7178-T6. Low power magnification, fractographs taken with a scanning electron, and optical microscopic examination of the milled sections revealed extensive pitting on both sides of the floors. Evidence found supports the conclusions that the floors failed by fatigue cracking that initiated near the center of the fuel-tank floor and ultimately propagated as rapid ductile-overload fractures. The fatigue cracks originated in pits on the fuel-cell side of the tank floors. The pits were attributed to attack caused by the alkaline-etch cleaning process. Recommendations included monitoring of the alkaline-etch cleaning to avoid the formation of pits and careful inspection following alkaline-etch cleaning, to be scheduled before release of the floor panels for painting.

Large quantities of aluminum-clad spent nuclear materials have been in interim storage in the fuel storage basins at The Savannah River Site while awaiting processing since 1989. This extended storage as a result of a moratorium on processing resulted in corrosion of the aluminum clad. Examinations of this fuel and other data from a corrosion surveillance program in the water basins have provided basic insight into the corrosion process and have resulted in improvements in the storage facilities and basin operations. Since these improvements were implemented, there has been no new initiation of pitting observed since 1993. This paper describes the corrosion of spent fuel and the metallographic examination of Mark 31A target slugs removed from the K-basin storage pool after 5 years of storage. It discusses the SRS Corrosion Surveillance Program and the improvements made to the storage facilities which have mitigated new corrosion in the basins.

Cracks occurred in a new ship hull after only three months in service. It was noted that the 5xxx series of aluminum alloys are often selected for weldability and are generally very resistant to corrosion. However, if the material has prolonged exposure at slightly elevated temperatures of 66 to 180 deg C (150 to 350 deg F), an alloy such as 5083 can become susceptible to intergranular corrosion. Investigation (visual inspection, corrosion testing, SEM images) supported the conclusion that the cracks occurred because during exposures to chloride solutions like seawater, galvanic couples formed between precipitates and the alloy matrix, leading to severe intergranular attack. 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.

Several pressurized air containers (i.e., diving tanks) made of non-heat-treatable Al-5Mg aluminum alloy failed catastrophically. Catastrophic failure occurred when a subcritical stress corrosion crack reached a critical size. Critical crack size for unstable propagation was reached prior to wall penetration, which could have led to subsequent loss of pressure, resulting in explosion of the cylinder. It was recommended that more stress corrosion resistant alloys be used for sea diving applications. Furthermore, cylinders should have a reduced wall thickness that can be determined employing the “leak-before-break” design philosophy, developed using fracture mechanics, to eliminate the possibility of catastrophic ruptures.

Failures of various types of hydraulic couplings used to connect pipes in a naval vessel are described and used to illustrate some of the general procedures for failure analysis. Cracking of couplings, which were manufactured from nickel-aluminum- bronze extruded bar, occurred in both seawater and air environments. Cracks initiated at an unusually wide variety of sites and propagated in either longitudinal or circumferential directions with respect to the axis of the couplings. Fracture surfaces were intergranular and exhibited little or no sign of corrosion (for couplings cracked in air), and there was very limited plasticity. Macroscopic progression markings were observed on fracture surfaces of several couplings but were not generally evident. At very high magnifications, numerous slip lines, progression markings, and striations were observed. In a few cases, where complete separation had occurred in service, small areas of dimpled overload fracture were observed. It was concluded from these observations, and from comparisons of cracks produced in service with cracks produced by laboratory testing under various conditions, that cracking had occurred by fatigue. The primary cause of failure was probably the unanticipated presence of high-frequency stress cycles with very low amplitudes, possibly due to vibration, resonance, or acoustic waves transmitted through the hydraulic fluid. Secondary causes of failure included the presence of high tensile residual stresses in one type of coupling, undue stress concentrations at some of the crack-initiation sites, and overtorquing of some couplings during installation. Recommendations on ways to prevent further failures based on these causes are discussed.

The Advanced Photon Source (APS) is a state-of-the-art synchrotron light source. The storage ring vacuum chamber is fabricated from 6061 extruded aluminum. Water connections to the vacuum chambers that were fabricated from 3003 aluminum had developed water leaks, which were subsequently remedied after considerable investigations.

Equipment in which an assembly of in-line cylindrical components rotated in water at 1040 rpm displayed excessive vibration after less than one hour of operation. The malfunction was traced to an aluminum alloy 6061-T6 combustion chamber that was part of the rotating assembly. Analysis (visual inspection, 100x/500x/800x micrographic examination, spectrographic analysis, and hardness testing) supported the conclusions that, as a result of improper heat treatment, the combustion-chamber material was too soft for successful use in this application. Misalignment of the combustion chamber and one or both of the mating parts resulted in eccentric rotation and the excessive vibration that caused malfunction of the assembly. Irregularities in the housing around the combustion chamber and temperature variation relating to the combustion pattern in the chamber were considered to be possible contributing factors to localization of the cavitation erosion. Recommendations included adopting inspection procedures to ensure that the specified properties of aluminum alloy 6061-T6 were obtained and that the combustion chamber and adjacent components were aligned within specified tolerances. In a similar situation, consideration should also be given to raising the pressure in the coolant in order to suppress the formation of cavitation bubbles.

Immediately after installation, leakage was observed at the mounting surface of several rebuilt hydraulic actuators that had been in storage for up to three years. At each joint, there was an aluminum alloy spacer and a vellum gasket. The mounting flanges of the steel actuators had been nickel plated. During assembly of the actuators a lubricant containing molybdenum disulfide had been applied to the gaskets as a sealant. The vellum gasket was found to be electrically conductive, and analysis (visual inspection, 500x unetched micrographs, galvanic action testing, and x-ray diffraction) supported the conclusions that leakage was the result of galvanic corrosion of the aluminum alloy spacers while in storage. The molybdenum disulfide was apparently suspended in a volatile water-containing vehicle that acted as an electrolyte between the aluminum alloy spacer and the nickel-plated steel actuator housing. Initially, the vellum gasket acted as an insulator, but the water-containing lubricant gradually impregnated the vellum gasket, establishing a galvanic couple. Recommendations included discontinuing use of molybdenum disulfide lubricant as a gasket sealer, and assembling the actuators using dry vellum gaskets.

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 forging of 7075-T6 aluminum alloy, which formed a support for the cylinder of a cargo door, cracked at an attachment hole. Fluorescent penetrant inspection showed the crack ran above and below the hole out onto the machined flat surface of the flange. A 6500x electron fractograph proved the crack to be a forging defect called a cold shut. Because defects of this type are usually detected when the raw forging is inspected, this occurrence was considered to be an isolated instance.