A turbine blade in an aircraft engine failed, fracturing at the root above the fir tree region. Fractography indicated that a fatigue crack initiated at the trailing edge of the blade and the final fracture occurred when the crack reached critical length. Although the exact cause of crack initiation could not be established, material defects, improper root loading, and high operating temperatures were ruled out. This chapter describes how investigators came to their conclusions and what they learned through visual and SEM examination and qualitative elemental analysis. It includes images of the microstructure and fracture surfaces and explains what some of the details reveal about the failure.

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.

A helicopter lost the outboard rib on a tail rotor blade in flight and was forced to land because of the resulting vibrations. The investigation that followed is described in this chapter along with key findings. As shown in a sketch, the rib is held in place by a set of six rivets. All of the rivets on the failed blade were missing and sections of skin were torn from most of the rivet holes. One such rivet hole was examined in a SEM, revealing corrosion on one of the tear surfaces and dimples (characteristic of ductile overload failure) on the other. In addition, the inner surface of the skin nearest the rib was found to be coated with soot, the paint on the leading edge of the top skin was abraded, and the skin in that area had thinned. Based on their findings, investigators concluded that the outboard rib separated because of stress-corrosion cracking around the rivets, and erosion may have contributed.

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.

A first-stage compressor blade failed prematurely in an aircraft engine, fracturing at the midpoint of the root transition region. An examination of the fracture surface revealed beach marks, striations, and pitting, indicating that the blade failed by fatigue due to a crack initiated by corrosion pits in the root transition region. The chapter recommends further investigations to determine the cause of pitting, which appears to be confined to the dovetail region.

The quill shaft in an aircraft engine was found in two pieces following a flameout. One piece was short, straight, and otherwise undamaged; the other piece was bent in several places as was the sleeve that covered it. The facture surface, as viewed under optical and scanning electron microscopes, was flat and shiny with deformation marks and dimples, typical of torsional overload, and signs of severe rubbing on the periphery. Based on their observations, investigators concluded that the quill shaft failed by torsional overload, the source of which could not be determined.

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.

An aircraft crashed following the loss of yaw control in full airborne flight. The subsequent discovery of broken shutter bolts in the rear pitch reaction control valve led to an inspection campaign that found bolt failures of a similar nature in valves on several other aircraft. The bolts were removed and analyzed to determine the mode and cause of failure. Based on the results of macroscopy, scanning electron fractography, metallographic examination, and chemical analysis, the failures were caused by stress corrosion cracking, and in one case, overtightening.

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.

A compressor blade made of titanium alloy fractured during an engine test. The material and processing conditions of the blade were found to be satisfactory, turning the focus of the investigation to operating anomalies and human error. A photograph of the failed blade shows well-defined chevron marks along the fracture surface that end in a shear lip on the convex side. Further examination using a SEM shows that the failure was due to overload. Based on these observations and the results of tensile testing and microstructural analysis, investigators concluded that a sudden impact load on the concave side of the blade caused it to fracture.

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.

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.

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.

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.

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.

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.

During a routine preflight inspection of a piston aircraft engine, part of a cooling fin was found that had broken off the cylinder. The piece, made of aluminum-silicon alloy, was cleaned and examined. Based on the fracture characteristics revealed by an electron microscope, it was concluded that the fin failed in a brittle manner by overload.

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.