Search Results for SEM fractography

This chapter discusses the failure of an aileron control cable in an aircraft and explains how investigators determined the cause. Based on their observations and the results of SEM fractography, investigators concluded that the cable had been damaged by a shearing tool, leading to its failure.

Fastening screws used in fuel-injection pumps failed during assembly and were examined to determine the cause. Based on observations and the result SEM fractography and hardness measurements, the screws failed by brittle intergranular fracture due to hydrogen embrittlement associated with plating procedures. The report includes recommendations for improving the quality of the screws.

This chapter discusses the failure of a control cable on an aircraft and the findings of an investigation that followed. The cable was made of stranded steel wire that was visibly worn. All seven strands had snapped and bore evidence of corrosion, pitting, nicks, and rubbing. Based on their observations and the results of SEM fractography, investigators concluded that tensile overload was the predominate cause of failure.

This chapter discusses the failure of a compressor blade in an aircraft engine and explains how investigators determined the cause. Based on visual examination and the results of SEM fractography and chemical analysis, it was concluded that blade failed due to fatigue fracture originating from nonmetallic inclusions in the blade root.

A second-stage turbine blade in an aircraft engine failed in service, fracturing along a path through the shroud hole. Cracks were also found in the shroud holes of the two adjacent blades. Based on the results of visual examination and SEM fractography, investigators concluded that the fracture and cracks were due to the fretting action of the pins inside the shroud holes.

After several failed attempts to lower their starboard wheels for landing, pilots engaged the help of gravity through g-force maneuvers and managed to coax the wheels into place. An inspection following the incident revealed a broken universal joint in one of the linkages that opens and closes the doors to the undercarriage compartment. The failed component was removed from the aircraft and examined using optical and electron microscopes. Under low magnification, the fracture surface appeared jagged except for one corner that was relatively smooth. SEM fractography revealed the presence of fatigue striations in the smooth region and dimpling elsewhere. Based on their findings, investigators concluded that fatigue loading initiated a crack in the universal joint that progressed with time and that the final fracture occurred due to bending tensile overload.

An aircraft went down over water some 30 minutes into a flight. The wreckage was retrieved and the elevator linkage components were dismantled, cleaned, and reassembled. As the chapter explains, both the port and starboard hinge pins had fractured at a tack welded joint along a flange. Based on visual examination, SEM fractography, and chemical analysis, investigators concluded that the hinge pins were not made from the specified steel and were not properly treated after cadmium plating. The pins failed due to hydrogen embrittlement, which may have been aggravated by welding. The chapter provides several recommendations to avoid such failures in the future.

Two compressor rotors of similar design and construction were severely damaged during operation. In one rotor, all the blades in the third and fourth stages had been sheared off and some had lifted from the dovetail portion of the drum. The damage in the other rotor was more extensive. Most of the blades in the first four stages had sheared off and many lifted from the dovetail region, particularly in the first two stages where several mounting dovetails had also fractured. Based on visual examination and the results of SEM fractography, metallography, and chemical analysis, investigators concluded that the compressor rotors failed due to stress-corrosion cracking in the dovetail mountings. They also provided recommendations to prevent or mitigate future occurrences.

A cardon shaft operating in an aircraft engine failed and was taken out and analyzed to determine the cause. A photograph of the broken shaft in the as-received condition shows the location and orientation of the fracture. The fracture surface appeared smooth, indicating that a considerable amount of rubbing occurred after the shaft broke. SEM fractography revealed deformation marks and elongated dimples, typical of shear overloads, along with other details. Based on their analysis, investigators concluded that the cardan shaft failed under torsional overload. They also cited a need for a more detailed examination of the driven end of the shaft.

A second-stage compressor blade in an aircraft engine fractured after 21 h of service. The remaining portion of the blade was removed and examined as were several adjacent blades. Based on the results of SEM fractography, microstructural analysis, and hardness testing, the blade failed due to stress-corrosion cracking combined with the effects of inadequate tempering.

A low-pressure turbine rotor blade failed during a test run, causing extensive damage to an aircraft engine. Visual examination showed that the nickel-base superalloy blade broke above the root platform in the airfoil section, leaving a fracture surface with two distinct regions, one characteristic of fatigue, the other, overload. Two dents were also visible on the leading edge, near the origin of the fracture. Based on these observations and the results of SEM fractography, investigators concluded that the blade failed due to fatigue aided by cracks in the surface coating caused by mechanical damage.

During a major servicing of an aircraft, cracks were found in the bottom wing root fitting. Based on dye penetrant inspection and the results of SEM fractography and chemical analysis, investigators concluded that the cracks were due to stress corrosion. They also recommended an inspection of all other aircraft with similar fittings and the consideration of alternate materials that are less prone to stress-corrosion cracking.

The aluminum alloy skin on the main rotor blade of a helicopter tore off in flight, and an investigation was subsequently conducted to find the cause. Visual examination and SEM fractography revealed that a fatigue crack originated on the underside of a rivet hole at the trailing edge of the blade. The crack then propagated through the outer skin toward the leading edge of the blade. Once the fatigue crack reached critical length, the sheet metal fractured catastrophically, tearing away from the blade.

Several components from the tail boom of a helicopter were found fractured at a crash site, including gusset plates, the hat section near the lower yoke, and a cable that controls the pitch of the tail rotor. The components were recovered from the wreckage and taken to a lab for closer examination. Based on their observations and the results of SEM fractography, failure analysts concluded that the gusset plates failed due to a downward bending overload in tension and that the tail rotor control cable snapped due to tensile overload. There were no indications of delayed failure in any of the areas examined.

Two filtration components installed on a developmental aircraft cracked during pressure impulse testing. Both parts were made from an aluminum alloy, solutionized and aged, and cracked due to fatigue. In both cases, the crack initiated at a transition region on an inner surface and progressed circumferentially outward. Based on these observations and the results of SEM fractography and microstructural analysis, the fatigue cracking can be traced to insufficient fillet radius at the transition zone.

During cyclic spin tests, the turbine disc in an aircraft engine broke apart with a loud noise, followed by a fire. Based on a detailed examination and the results of SEM fractography and hardness measurements, failure analysts concluded that a locking plate became dislodged due to the shearing of the screws that hold it in place. They also provided recommendations to remediate the problem.

The nose landing gear on an aircraft malfunctioned during landing roll. After the incident, two fractured studs were found in the retraction jack support beam. Based on visual examination and the results of SEM fractography, investigators concluded that the studs failed by fatigue, a vulnerability because of the way they were mounted. A sketch showing the correct mounting configuration is included in the report.

A high-pressure turbine blade in an aircraft engine failed prematurely, fracturing close to the root. Visual examination revealed significant plastic deformation on the leading edge of the blade, blocky cleavage on the trailing edge, and a region covered with fissures in between. Based on their observations and the results of SEM imaging described in the chapter, investigators concluded that the blade failed by low-cycle fatigue, acting on a preexisting crack.

The tail rotor blade of a helicopter developed a visible crack during service. The cracked region was removed from the blade and the fracture surface was examined in a SEM, revealing shallow pitting on the inside surface of the skin and a corresponding reduction in thickness. Based on these findings, investigators concluded that the failure was due to a fatigue crack initiated from a corrosion pit, which may have been caused by chemicals released by the burning of bonding resin.