Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0046142
To ensure no malfunctions and although there were no apparent problems, a main fuel control was returned to the factory for examination after service on a test aircraft engine that had experienced high vibrations. When the fuel control was disassembled, a lever, cast from AMS 5350 (AISI type 410) stainless steel that was through-hardened to 26 to 32 HRC and passivated, was shown to be cracked. The crack initiated at the sharp corner of the elongated milled slot and propagated across to the outer wall. The sections around the crack were spread about 30 deg apart, showing the fracture surface under investigation had beach marks initiating at the sharp corner along the milled slot. Changes in frequency or amplitude of vibration caused different rates of propagation, resulting in a change in pattern. This evidence supported the conclusion that the lever failed in fatigue as a result of excessive vibration of the fuel control on the test engine. Recommendations included redesign of the lever with a large radius in the corner where cracking originated. This would reduce the stress-concentration factor significantly, thus minimizing the susceptibility of the lever to fatigue.
{"title":"Fatigue Fracture of a Cast Stainless Steel Lever","authors":"","doi":"10.31399/asm.fach.aero.c0046142","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0046142","url":null,"abstract":"\u0000 To ensure no malfunctions and although there were no apparent problems, a main fuel control was returned to the factory for examination after service on a test aircraft engine that had experienced high vibrations. When the fuel control was disassembled, a lever, cast from AMS 5350 (AISI type 410) stainless steel that was through-hardened to 26 to 32 HRC and passivated, was shown to be cracked. The crack initiated at the sharp corner of the elongated milled slot and propagated across to the outer wall. The sections around the crack were spread about 30 deg apart, showing the fracture surface under investigation had beach marks initiating at the sharp corner along the milled slot. Changes in frequency or amplitude of vibration caused different rates of propagation, resulting in a change in pattern. This evidence supported the conclusion that the lever failed in fatigue as a result of excessive vibration of the fuel control on the test engine. Recommendations included redesign of the lever with a large radius in the corner where cracking originated. This would reduce the stress-concentration factor significantly, thus minimizing the susceptibility of the lever to fatigue.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"47 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129268089","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0090030
A DC-10 in transit from Denver to Chicago experienced failure of the center engine. The titanium compressor disk burst and severed the hydraulics of the plane. Investigation supports the conclusion that the cause of the disk rupture was the presence of a large fatigue crack near the bore emanating from a hard alpha (HA) defect. Such defects can result from occasional upsets during the vacuum melting of titanium. These nitrogen-rich alpha titanium anomalies are brittle and often have associated microcracks and microvoids. A probabilistic damage tolerance approach was recommended to address the anomalies, with the objective of enhancing rotor life management practices. The ongoing work involves the use of fracture mechanics and software (called DARWIN.) optimized for damage tolerant design and analysis of metallic structural components.
{"title":"Probabilistic Damage Tolerance Analysis of Gas Turbine Rotors","authors":"","doi":"10.31399/asm.fach.aero.c0090030","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0090030","url":null,"abstract":"\u0000 A DC-10 in transit from Denver to Chicago experienced failure of the center engine. The titanium compressor disk burst and severed the hydraulics of the plane. Investigation supports the conclusion that the cause of the disk rupture was the presence of a large fatigue crack near the bore emanating from a hard alpha (HA) defect. Such defects can result from occasional upsets during the vacuum melting of titanium. These nitrogen-rich alpha titanium anomalies are brittle and often have associated microcracks and microvoids. A probabilistic damage tolerance approach was recommended to address the anomalies, with the objective of enhancing rotor life management practices. The ongoing work involves the use of fracture mechanics and software (called DARWIN.) optimized for damage tolerant design and analysis of metallic structural components.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"3 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126888357","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0006394
A preflight inspection found a broken diaphragm from a side controller fabricated from 17-7 PH stainless steel in the RH 950 heat treatment condition. Failure occurred by cracking of the base of the flange-like diaphragm. The crack traveled 360 deg around the diaphragm. Scanning electron microscopy (SEM) revealed that the failure occurred by a brittle intergranular mechanism and stress-corrosion cracking (SCC), and indicated a failure mode of selective grain-boundary separation. The diaphragms were heat treated in batches of 25. An improper heat treatment could have resulted in the formation of grain boundary precipitates, including chromium carbides. It was concluded that failure of the diaphragm was due to a combination of sensitization caused by improper heat treatment and subsequent SCC. It was recommended that the remaining 24 sensor diaphragms from the affected batch be removed from service. In addition, a sample from each heat treat batch should be submitted to the Strauss test (ASTM A262, practice E) to determine susceptibility to intergranular corrosion. Also, it was recommended that a stress analysis be performed on the system to determine whether a different heat treatment (which would offer lower strength but higher toughness) could be used for this part.
{"title":"Failure of an Aircraft Controller Diaphragm","authors":"","doi":"10.31399/asm.fach.aero.c0006394","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0006394","url":null,"abstract":"\u0000 A preflight inspection found a broken diaphragm from a side controller fabricated from 17-7 PH stainless steel in the RH 950 heat treatment condition. Failure occurred by cracking of the base of the flange-like diaphragm. The crack traveled 360 deg around the diaphragm. Scanning electron microscopy (SEM) revealed that the failure occurred by a brittle intergranular mechanism and stress-corrosion cracking (SCC), and indicated a failure mode of selective grain-boundary separation. The diaphragms were heat treated in batches of 25. An improper heat treatment could have resulted in the formation of grain boundary precipitates, including chromium carbides. It was concluded that failure of the diaphragm was due to a combination of sensitization caused by improper heat treatment and subsequent SCC. It was recommended that the remaining 24 sensor diaphragms from the affected batch be removed from service. In addition, a sample from each heat treat batch should be submitted to the Strauss test (ASTM A262, practice E) to determine susceptibility to intergranular corrosion. Also, it was recommended that a stress analysis be performed on the system to determine whether a different heat treatment (which would offer lower strength but higher toughness) could be used for this part.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"51 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129468076","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0047125
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.
{"title":"Fatigue Cracking That Originated at a Material Defect in a Forged Aircraft Wheel Half","authors":"","doi":"10.31399/asm.fach.aero.c0047125","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0047125","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"12 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125339537","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c9001907
M. Pepi
Aircraft missile launcher attachment bolts fabricated from cadmium-coated Hy-tuf steel were found broken. Subsequent analysis of the broken bolts indicated three causes of failure. First, the bolts had been carburized, which was not in conformance with the heat treating requirements. Second, macroetching showed that the bolts has been machined from stock rather than forged, and the threads cut rather than rolled. It was also determined that hydrogen-assisted stress-corrosion cracking also played a part in the failure of the high-strength bolts.
{"title":"Failure of Steel Missile Launcher Attachment Bolts","authors":"M. Pepi","doi":"10.31399/asm.fach.aero.c9001907","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001907","url":null,"abstract":"\u0000 Aircraft missile launcher attachment bolts fabricated from cadmium-coated Hy-tuf steel were found broken. Subsequent analysis of the broken bolts indicated three causes of failure. First, the bolts had been carburized, which was not in conformance with the heat treating requirements. Second, macroetching showed that the bolts has been machined from stock rather than forged, and the threads cut rather than rolled. It was also determined that hydrogen-assisted stress-corrosion cracking also played a part in the failure of the high-strength bolts.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"4 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125402074","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0091678
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.
{"title":"Stress-Corrosion Cracking of a Forged Aircraft Lug","authors":"","doi":"10.31399/asm.fach.aero.c0091678","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0091678","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"29 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127775971","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0006436
A failed H-11 tool steel pylon attachment stud was found during a routine walk-around inspection. The stud exhibited gross localized corrosion pitting at several different areas on its surface. Light general rust was also evident. Severe pitting occurred near the fracture location. The fracture face contained evidence of intergranular SCC as well as ductile dimples. The protective coating was found to be an inorganic water-base aluminide coating having a coating thickness of 7.5 to 13 micron (0.3 to 0.5 mil). The coating was of a nonuniform mottled nature. It was concluded that the failure of the pylon attachment stud was caused by general corrosion followed by SCC. The stud was not adequately protected against corrosion by the coating. It was recommended that the coating be applied to a thickness of 38 to 75 micron (1.5 to 3 mil) to provide long-time corrosion resistance. The coating must be either burnished or cured at 540 deg C (1000 deg F) to provide cathodic protection to the steel. Other coatings, such as cadmium or aluminum, were also recommended if a thinner coating is needed.
{"title":"Failure of Tool Steel Pylon Attachment Stud","authors":"","doi":"10.31399/asm.fach.aero.c0006436","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0006436","url":null,"abstract":"\u0000 A failed H-11 tool steel pylon attachment stud was found during a routine walk-around inspection. The stud exhibited gross localized corrosion pitting at several different areas on its surface. Light general rust was also evident. Severe pitting occurred near the fracture location. The fracture face contained evidence of intergranular SCC as well as ductile dimples. The protective coating was found to be an inorganic water-base aluminide coating having a coating thickness of 7.5 to 13 micron (0.3 to 0.5 mil). The coating was of a nonuniform mottled nature. It was concluded that the failure of the pylon attachment stud was caused by general corrosion followed by SCC. The stud was not adequately protected against corrosion by the coating. It was recommended that the coating be applied to a thickness of 38 to 75 micron (1.5 to 3 mil) to provide long-time corrosion resistance. The coating must be either burnished or cured at 540 deg C (1000 deg F) to provide cathodic protection to the steel. Other coatings, such as cadmium or aluminum, were also recommended if a thinner coating is needed.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"437 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122145475","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0046146
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.
{"title":"Fatigue Fracture of an Aluminum Alloy 7075-T73 Landing-Gear Torque Arm","authors":"","doi":"10.31399/asm.fach.aero.c0046146","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0046146","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127459223","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c9001529
L. M. Gammon, M. Hyatt, G. H. Narayanan, H. Oberson, Harcayal B. Singh
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.
{"title":"Analysis of Aircraft Damage from a Lightning Strike","authors":"L. M. Gammon, M. Hyatt, G. H. Narayanan, H. Oberson, Harcayal B. Singh","doi":"10.31399/asm.fach.aero.c9001529","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001529","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"38 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134095598","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c0006398
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.
{"title":"Fuel Line Corrosion","authors":"","doi":"10.31399/asm.fach.aero.c0006398","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0006398","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"10 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129383092","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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