A group of control valves that regulate production in a field of sour gas wellheads performed satisfactorily for three years before pits and cracks were detected during an inspection. One of the valves was examined using chemical and microstructural analysis to determine the cause of failure and provide preventive measures. The valve body was made of A216-WCC cast carbon steel. Its inner surface was covered with cracks stemming from surface pits. Investigators concluded that the failure was caused by a combination of hydrogen-induced corrosion cracking and sulfide stress-corrosion cracking. Based on test data and cost, A217-WC9 cast Cr–Mo steel would be a better alloy for the application.

A hollow, splined alloy steel aircraft shaft (machined from an AMS 6415 steel forging – approximately the same composition as 4340 steel – then quenched and tempered to a hardness of 44.5 to 49 HRC) cracked in service after more than 10,000 h of flight time. The inner surface of the hollow shaft was exposed to hydraulic oil at temperatures of 0 to 80 deg C (30 to 180 deg F). Analysis (visual inspection, 15-30x low magnification examination, 4x light fractograph, chemical analysis, hardness testing) supported the conclusions that the shaft cracked in a region subjected to severe static radial, cyclic torsional, and cyclic bending loads. Cracking originated at corrosion pits on the smoothly finished surface and propagated as multiple small corrosion-fatigue cracks from separate nuclei. The originally noncorrosive environment (hydraulic oil) became corrosive in service because of the introduction of water into the oil. Recommendations included taking additional precautions in operation and maintenance to prevent the use of oil containing any water through filling spouts or air vents. Also, polishing to remove pitting corrosion (but staying within specified dimensional tolerances) was recommended as a standard maintenance procedure for shafts with long service lives.

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.

Military aircraft use a cartridge ignition system for emergency engine starts. Analysis of premature failures of steel (AISI 4340) breech chambers in which the solid propellant cartridges were burned identified corrosion as one problem with an indication that stress-corrosion cracking may have occurred. A study was made for stress-corrosion cracking susceptibility of 4340 steel in a paste made of the residues collected from used breech chambers. The constant extension rate test (CERT) technique was employed and SCC susceptibility was demonstrated. The residues, which contained both combustion products from the cartridges and corrosion products from the chamber, were analyzed using elemental analysis and x-ray diffraction techniques. Electrochemical polarization techniques were also utilized to estimate corrosion rates.

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.

A T-bolt was part of the coupling for a bleed air duct of a jet engine on a transport plane. Specifications required that the 4.8 mm diam component be made of AISI type 431 stainless steel and heat treated to 44 HRC. The operating temperature of the duct is 425 to 540 deg C (800 to 1000 deg F), but that of the bolt is lower. The T-bolt broke after three years of service. The expected service life was equal to that of the aircraft. It was found that the bolt broke as a result of SCC. Thermal stresses were induced into the bolt by intermittent operation of the jet engine. Mechanical stresses were induced by tightening of the clamp around the duct, which in effect acted to straighten the bolt. The action of these stresses on the carbides that precipitated in the grain boundaries resulted in fracture of the bolt. Due to the operating temperatures of the duct near the bolt, the material was changed to A-286, which is less susceptible to carbide precipitation. The bolt is strengthened by shot peening and rolling the threads after heat treatment. Avoiding temperatures in the sensitizing range is desirable, but difficult to ensure because of the application.

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.

Examination of a cracked nose landing gear cylinder made of AISI 4340 Cr-Mo-Ni alloy steel proved that the part started to fail on the inside diam. When the nucleus of the stress-corrosion crack was studied in detail, iron oxide was found on the fracture surface. A 6500x micrograph revealed this area also displayed an intergranular texture. One of a group of small grinding cracks on the ID of the cylinder nucleated the failure. Other evidence indicated the cracks developed when the cylinder was ground during overhaul.

Ground maintenance personnel discovered hydraulic fluid leaking from two small cracks in a main landing gear cylinder made from AISI 4340 Cr-Mo-Ni alloy steel. Failure of the part had initiated on the ID of the cylinder. Numerous cracks were found under the chromium plate. A 6500x electron fractograph showed cracking was predominantly intergranular with hairline indications. Leaking had occurred only 43 h after overhaul of the part. Total service time on the part was 9488 h. It was concluded that cracking on the ID was caused by hydrogen embrittlement which occurred during or after overhaul. The specific source of hydrogen which produced failure was not ascertainable.

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 nose landing gear cylinder made from AISI 4340 Ni-Cr-Mo alloy steel was found cracked and leaking, causing partial depressurization. Investigation revealed the crack to be a stress-corrosion type, judging by the 6500x electron fractograph. It had started in a region of concentrated, large non-metallic inclusions near the chromium-plated ID of the cylinder. Also, there were breaks in the chromium plate and pits in the underlying base metal. The cylinder had been in service for 18,017 h, and 5948 h had passed since the first and only overhaul. Substandard plating of the ID at this time ultimately resulted in pitting of the metal. The combination of surface pitting and stress concentration at the nearby inclusions resulted in stress-corrosion cracking.

A commercial aircraft wheel half, machined from an aluminum alloy 2014 forging that had been heat treated to the T6 temper, was removed from service because a crack was discovered in the area of the grease-dam radius during a routine inspection. Neither the total number of landings nor the roll mileage was reported, but about 300 days had elapsed between the date of manufacture and the date the wheel was removed from service. The analysis (visual inspection, macrographs, micrographs, electron microprobe) supported the conclusions that the wheel half failed by fatigue. The fatigue crack originated at a material imperfection and progressed in more than one plane because changes in the direction of wheel rotation altered the direction of the applied stresses. Recommendations included rewriting the inspection specifications to require sound forgings.

Pitostatic system connectors were being found cracked on several aircraft. Two of the cracked connectors made of 2024-T351 aluminum alloy were submitted for failure analysis. The connectors had cut pipelike threads that were sealed with Teflon-type tape when installed. Longitudinal cracks were located near the opening of the female ends of each connector. A cross section showed intergranular cracking with multiple branching in one connector. Scanning electron microscopy (SEM) showed intergranular cracking and separation of elongated grains. A cross section of connector threads showed an incomplete thread form resulting from improper tapping. It was concluded that the pitostatic system connectors failed by SCC. The stress was caused by forcing the improperly threaded female nut over its fully threaded male counterpart to effect a seal. The one connector tested for chemical composition was not made of 2024 aluminum alloy as reported but of 2017 aluminum. It was recommended that the pitostatic system connector manufacturing process be revised to produce full-depth threads rather than pseudo pipe threads. Wall thickness should be increased to increase the hoop stress bearing area if pipe threads were to be used. A determination of proper torque values for tightening the connectors was suggested also.

Examination of a 7075-T6 aluminum alloy pylon strut revealed cracks in two locations on the ears of the strut. Because the part was still intact, the cracks had to be forced open so that the fractures could be examined. Scanning electron microscopy (SEM) of the opened cracks showed that the crack surfaces were covered with a mud crack pattern suggestive of stress-corrosion cracking (SCC). The T6 temper is susceptible to SCC. It was concluded that cracking of the strut could have been aggravated by the hard landing experienced by the aircraft. The strut, however, contained stress-corrosion cracks which were present before the landing. It was recommended that an inspection for SCC be made of all pylon struts with a similar service life.

A welded elbow assembly (AISI type 321 stainless steel, with components joined with ER347 stainless steel filler metal by gas tungsten arc welding) was part of a hydraulic-pump pressure line for a jet aircraft. The other end of the tube was attached to a flexible metal hose, which provided no support and offered no resistance to vibration. The line was leaking hydraulic fluid at the nut end of the elbow. Investigation supported the conclusion that failure was by fatigue cracking initiated from a notch at the root of the weld and was propagated by cyclic loading of the tubing as the result of vibration and inadequate support of the hose assembly. Recommendations included changing the joint design from a cylindrical lap joint to a square-groove butt joint. Also, an additional support was recommended for the hose assembly to minimize vibration at the elbow.

The master connecting rod of a reciprocating aircraft engine revealed cracks during routine inspection. The rods were forged from 4337 (AMS 6412) steel and heat treated to a specified hardness of 36 to 40 HRC. H-shaped cracks in the wall between the knuckle-pin flanges were revealed by visual examination. The cracks were originated as circumferential cracks and then propagated transversely into the bearing-bore wall. No inclusions in the master rod were detected by magnetic-particle and x-ray inspection. Three large inclusions lying approximately parallel to the grain direction and fatigue beach marks around two of the inclusions were revealed by macroscopic examination of the fracture surface. Large nonmetallic inclusions that consisted of heavy concentrations of aluminum oxide (Al2O3) were revealed by microscopic examination of a section through the fracture origin. The forging vendors were notified about the excess size of the nonmetallic inclusions in the master connecting rods and a nondestructive-testing procedure for detection of large nonmetallic inclusions was established.

An aircraft fuel-nozzle-support assembly exhibited cracks along the periphery of a fusion weld that attached a support arm to a fairing in a joint that approximated a T-shape in cross section. The base metal was type 321 stainless steel. Examination showed a good-quality weld penetrating to the support arm beneath, but revealed notch configurations at the inner mating surfaces at each edge of the fairing, the result of welding a poor fit-up of the support arm to the fairing. Fractures that originated at the cracks were examined by stereomicroscope and were found to contain fatigue marks that indicated crack propagation from multiple origins at the inner surface of the weld edge. Fatigue cracking was initiated at stress concentrations created by the notches at the inner surfaces between the support arm and the fairing, enhanced by poor fit-up in preparation for welding.

A 2000-T6 aluminum alloy bracket failed in a coastal environment because corrosive chlorides got between the bracket and attachment bolt. The material used for the part was susceptible to stress corrosion under the service conditions. Cracking may have been aggravated by galvanic action between aluminum alloy bracket and steel bolt. To preclude or minimize recurrences, fittings in service should be inspected periodically by dye penetrant for signs of cracking on the end face and within the fitting hole and protected with a suitable coating to exclude damaging chlorides. Also, a 2000-T6 aluminum alloy swivel fitting experienced intergranular corrosion fracture as the result of stress-accelerated corrosion. Corrosion began because of a loose fit between the aluminum swivel fitting and steel tube assembly, which caused fretting. Inadequate maintenance and/or abnormal service operation may have loosened the fitting.

A liquid hydrogen main fuel control valve for a rocket engine failed by fracture of the Ti-5Al-2.5Sn body during the last of a series of static engine test firings. Fractographic, metallurgical, and stress analyses determined that a combination of fatigue and unexpected aqueous stress-corrosion cracking initiated and propagated the crack which caused failure. The failure analysis approach and its results are described to illustrate how fractography and fracture mechanics, together with a knowledge of the crack initiation and propagation mechanisms of the valve material under various stress states and environments, helped investigators to trace the cause of failure.

A steam-condensate line (type 316 stainless steel tubing) began leaking after five to six years in service. The line carried steam condensate at 120 deg C (250 deg F) with a two hour heat-up/cool-down cycle. No chemical treatment had been given to either the condensate or the boiler water. To check for chlorides, the inside of the tubing was rinsed with distilled water, and the rinse water was collected in a clean beaker. A few drops of silver nitrate solution were added to the rinse water, which clouded slightly because of the formation of insoluble silver chloride. This and additional investigation (visual inspection, and 250x micrograph etched with aqua regia) supported the conclusion that the tubing failed by chloride SCC. Chlorides in the steam condensate also caused corrosion of the inner surface of the tubing. Stress was produced when the tubing was bent during installation. Recommendations included providing water treatment to remove chlorides from the system. Continuous flow should be maintained throughout the entire tubing system to prevent concentration of chlorides. No chloride-containing water should be permitted to remain in the system during shutdown periods, and bending of tubing during installation should be avoided to reduce residual stress.