Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c9001903
Jivan B. Shah
Fretting and/or fretting corrosion fatigue have been observed on such parts as main rotor counterweight tie rods, fixed-pitch propeller blades, propeller blade clamps, pressure regulator lines, and landing gear support brackets. Microcracks started from severe corrosion pits in a failed control rotor spar tube assembly made of cadmium-plated AISI 4130 Cr-Mo alloy steel. Inadequate design was responsible for the failure. A lower tine of the main rotor blade cuff failed in fatigue. The rotor blade cuff was forged of 2014-T6 aluminum alloy. Initial stages of crack growth displayed features typical of low stress intensity fatigue of aluminum alloys. The fatigue resulted from abnormal fretting owing to inadequate torquing of the main retention bolts. Aircraft maintenance engineers and owners were advised to adhere to specifications when torquing this joint.
{"title":"Fretting and Corrosion in Aircraft Components","authors":"Jivan B. Shah","doi":"10.31399/asm.fach.aero.c9001903","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001903","url":null,"abstract":"\u0000 Fretting and/or fretting corrosion fatigue have been observed on such parts as main rotor counterweight tie rods, fixed-pitch propeller blades, propeller blade clamps, pressure regulator lines, and landing gear support brackets. Microcracks started from severe corrosion pits in a failed control rotor spar tube assembly made of cadmium-plated AISI 4130 Cr-Mo alloy steel. Inadequate design was responsible for the failure. A lower tine of the main rotor blade cuff failed in fatigue. The rotor blade cuff was forged of 2014-T6 aluminum alloy. Initial stages of crack growth displayed features typical of low stress intensity fatigue of aluminum alloys. The fatigue resulted from abnormal fretting owing to inadequate torquing of the main retention bolts. Aircraft maintenance engineers and owners were advised to adhere to specifications when torquing this joint.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"54 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":"116451775","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.c9001540
J. A. Scott
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.
{"title":"Use of Electron Fractography to Diagnose Hairline Cracking in an AISI 4340 Main Landing Gear Cylinder","authors":"J. A. Scott","doi":"10.31399/asm.fach.aero.c9001540","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001540","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"157 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":"116603731","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.c9001575
L.D. Cologna
An oil scavenge pump was found to have failed when a protective shear neck fractured during the start of a jet engine. Visual inspection revealed that the driven gear in one of the bearing compartments was frozen as was the corresponding drive gear. Spacer wear and thermal discoloration (particularly on the driven gear) were also observed. The gears were made from 32Cr-Mo-V13 steel, hardened and nitrided to 750 to 950 HV. Micrographic inspection of the gear teeth revealed microstructural changes that, in context, appear to be the result of friction heating. The spacers consist of Cu alloy (AMS4845) bushings force fit into AA2024-T3 Al alloy spacing elements. It was found that uncontrolled fit interference between the two components had led to Cu alloy overstress. Thermal cycling under operating conditions yielded the material. The dilation was directed inward to the shaft, however, because the bushing had only a few microns of clearance. The effect caused the oil to squeeze out, resulting in metal-to-metal contact, and ultimately failure.
{"title":"Investigation of Jet Engine Oil Scavenge Pump Failure","authors":"L.D. Cologna","doi":"10.31399/asm.fach.aero.c9001575","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001575","url":null,"abstract":"\u0000 An oil scavenge pump was found to have failed when a protective shear neck fractured during the start of a jet engine. Visual inspection revealed that the driven gear in one of the bearing compartments was frozen as was the corresponding drive gear. Spacer wear and thermal discoloration (particularly on the driven gear) were also observed. The gears were made from 32Cr-Mo-V13 steel, hardened and nitrided to 750 to 950 HV. Micrographic inspection of the gear teeth revealed microstructural changes that, in context, appear to be the result of friction heating. The spacers consist of Cu alloy (AMS4845) bushings force fit into AA2024-T3 Al alloy spacing elements. It was found that uncontrolled fit interference between the two components had led to Cu alloy overstress. Thermal cycling under operating conditions yielded the material. The dilation was directed inward to the shaft, however, because the bushing had only a few microns of clearance. The effect caused the oil to squeeze out, resulting in metal-to-metal contact, and ultimately failure.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"36 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":"117074051","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.c9001745
Jivan B. Shah
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.
{"title":"Corrosion Cracking of Aircraft Components","authors":"Jivan B. Shah","doi":"10.31399/asm.fach.aero.c9001745","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001745","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"54 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":"114660626","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.c0091809
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.
{"title":"Microbially Influenced Corrosion of Aluminum Alloy Aircraft Fuel Tanks","authors":"","doi":"10.31399/asm.fach.aero.c0091809","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0091809","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"42 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":"127918378","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.c0047072
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.
{"title":"Fatigue Failure of an Aluminum Alloy Assembly at Spot Welds Because of Improper Heat Treatment","authors":"","doi":"10.31399/asm.fach.aero.c0047072","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0047072","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"22 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":"123951283","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.c9001491
M. Katcher
Two silica phenolic nozzle liners cracked during proof testing. The test consisted of pressuring the nozzles to 14.1 MPa (2050 psia) for 5 to 20 s. It was concluded that the failure was due to longitudinal cracking in the convergent exhaust-nozzle insulators, stemming from the use of silica phenolic tape produced from flawed materials that went undetected by the quality control tests, which at the time, assessed tape strength properties in the warp rather than the bias direction. Once the nozzle manufacturer and its suppliers identified the problem, they changed their quality control procedures and resumed production of nozzle liners with more tightly controlled fiber/fabric materials.
{"title":"Failure Analysis of Silica Phenolic Nozzle Liners","authors":"M. Katcher","doi":"10.31399/asm.fach.aero.c9001491","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001491","url":null,"abstract":"\u0000 Two silica phenolic nozzle liners cracked during proof testing. The test consisted of pressuring the nozzles to 14.1 MPa (2050 psia) for 5 to 20 s. It was concluded that the failure was due to longitudinal cracking in the convergent exhaust-nozzle insulators, stemming from the use of silica phenolic tape produced from flawed materials that went undetected by the quality control tests, which at the time, assessed tape strength properties in the warp rather than the bias direction. Once the nozzle manufacturer and its suppliers identified the problem, they changed their quality control procedures and resumed production of nozzle liners with more tightly controlled fiber/fabric materials.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"9 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":"128077097","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.c9001558
D. Meyn, R. Bayles
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.
{"title":"Analysis of Sub-Critical Cracking in a Ti-5Al-2.5Sn Liquid Hydrogen Control Valve","authors":"D. Meyn, R. Bayles","doi":"10.31399/asm.fach.aero.c9001558","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001558","url":null,"abstract":"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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"95 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":"125177047","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.c9001018
C. Howard Craft
Initial investigation showed that a landing gear failure was the result of a hard landing with no evidence of contributory factors. The objective of reexamination was to determine whether there was any evidence of metallurgical failure. The landing gear was primarily an AISI type 6150 Cr-V steel flat spring attached at the top end to the fuselage and at the bottom end to the axle. Failure occurred at the clamping point near the top end of this spring. The failure showed evidence of severe brinelling at one corner in the clamping area. The fracture surfaces were clean, fresh, and indicative of a shock type of failure pattern. Closer examination, however, showed a fatigue crack at one corner. At this point, there was definite evidence of progression and oxidation. It was concluded that the corner in question was subjected to repeated brinelling resulting from normal landing loads, probably accentuated by looseness in the clamping device. The resulting residual tensile stress lowered the effective fatigue strength at that point against drag and side loads.
{"title":"Failure of a Main Landing Gear on a Light Airplane","authors":"C. Howard Craft","doi":"10.31399/asm.fach.aero.c9001018","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001018","url":null,"abstract":"\u0000 Initial investigation showed that a landing gear failure was the result of a hard landing with no evidence of contributory factors. The objective of reexamination was to determine whether there was any evidence of metallurgical failure. The landing gear was primarily an AISI type 6150 Cr-V steel flat spring attached at the top end to the fuselage and at the bottom end to the axle. Failure occurred at the clamping point near the top end of this spring. The failure showed evidence of severe brinelling at one corner in the clamping area. The fracture surfaces were clean, fresh, and indicative of a shock type of failure pattern. Closer examination, however, showed a fatigue crack at one corner. At this point, there was definite evidence of progression and oxidation. It was concluded that the corner in question was subjected to repeated brinelling resulting from normal landing loads, probably accentuated by looseness in the clamping device. The resulting residual tensile stress lowered the effective fatigue strength at that point against drag and side loads.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"86 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":"124971126","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.c9001747
Jivan B. Shah
This report covers case histories of failures in fixed-wing light airplane and helicopter components. In a 2025-T6 or 2219 aluminum alloy propeller blade that failed near the tip, cracks started on the leading edge at surface damage in the critical area-the zone between 4 and 10 in. from the tip of the blade. Incorrect dressing and inadequate pre-flight inspection were the two main causes. Two other types of propeller blade fatigue failures resulted mainly from propeller straightening operations, usually performed after previous blade bending damage. To eliminate blade tip failures, all surface-damaged material should be removed and polished smooth before further flight. The blade should be correctly dressed. Also, the tachometer should be calibrated to ensure the engine/propeller combination is not operated in the critical speed range at normal cruising speeds.
{"title":"Damaged Propeller Blades","authors":"Jivan B. Shah","doi":"10.31399/asm.fach.aero.c9001747","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001747","url":null,"abstract":"\u0000 This report covers case histories of failures in fixed-wing light airplane and helicopter components. In a 2025-T6 or 2219 aluminum alloy propeller blade that failed near the tip, cracks started on the leading edge at surface damage in the critical area-the zone between 4 and 10 in. from the tip of the blade. Incorrect dressing and inadequate pre-flight inspection were the two main causes. Two other types of propeller blade fatigue failures resulted mainly from propeller straightening operations, usually performed after previous blade bending damage. To eliminate blade tip failures, all surface-damaged material should be removed and polished smooth before further flight. The blade should be correctly dressed. Also, the tachometer should be calibrated to ensure the engine/propeller combination is not operated in the critical speed range at normal cruising speeds.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"2011 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":"131808406","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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