Pub Date : 2019-06-01DOI: 10.31399/asm.fach.aero.c9001505
J. Hutchinson
A helicopter had just taken off when there was a loud bang and the engine started to overspeed. After landing and inspection, the transmission was disassembled. It was discovered that the assembly containing the output shaft to the main rotor had failed. The output shaft assembly was made up of two parts: the output shaft with an integral 10 in. diam upper disc at approximately mid-section; and a 10 in. diam lower disc. During manufacture, the lower disc was attached to the output shaft by an electron beam weld. The fracture had a single fatigue initiation site, coincident with the annular zone of remelted material on the inner surface of the disc. In the lower disc, the fracture was also 80% fatigue, but high stress, low cycle in nature and contained multiple initiation sites coincident with an electron beam weld bead. It was concluded that fatigue in the upper disc resulted from the presence of a metallurgical stress concentration caused by the electron weld beam impingement on the inner surface of the upper disc. An Airworthiness Directive was issued, and the manufacturer issued a mandatory service bulletin outlining a periodic inspection for the output shaft assembly.
{"title":"Helicopter Main Rotor Transmission Output Shaft Failure","authors":"J. Hutchinson","doi":"10.31399/asm.fach.aero.c9001505","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001505","url":null,"abstract":"\u0000 A helicopter had just taken off when there was a loud bang and the engine started to overspeed. After landing and inspection, the transmission was disassembled. It was discovered that the assembly containing the output shaft to the main rotor had failed. The output shaft assembly was made up of two parts: the output shaft with an integral 10 in. diam upper disc at approximately mid-section; and a 10 in. diam lower disc. During manufacture, the lower disc was attached to the output shaft by an electron beam weld. The fracture had a single fatigue initiation site, coincident with the annular zone of remelted material on the inner surface of the disc. In the lower disc, the fracture was also 80% fatigue, but high stress, low cycle in nature and contained multiple initiation sites coincident with an electron beam weld bead. It was concluded that fatigue in the upper disc resulted from the presence of a metallurgical stress concentration caused by the electron weld beam impingement on the inner surface of the upper disc. An Airworthiness Directive was issued, and the manufacturer issued a mandatory service bulletin outlining a periodic inspection for the output shaft assembly.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"20 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":"114653068","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.c0046870
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.
{"title":"Corrosion-Fatigue Cracking in an AMS 6415 Steel Aircraft Shaft","authors":"","doi":"10.31399/asm.fach.aero.c0046870","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0046870","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"5 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":"125124559","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.c9001906
M. Pepi
Suspension lugs fabricated from AISI 4340 steel used to facilitate loading of bombs onto the underside of military aircraft could not sustain required loads during routine proof load testing. Three failed lugs underwent visual examination, chemical analysis, metallography, hardness testing, scanning electron microscopy, and energy-dispersive x-ray spectroscopy. It was determined that the failures were due to forging defects. Both forging laps and seams acted as stress concentrators when the lugs were loaded during proof testing.
{"title":"Failure of a Steel Bomb Suspension Lug","authors":"M. Pepi","doi":"10.31399/asm.fach.aero.c9001906","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001906","url":null,"abstract":"\u0000 Suspension lugs fabricated from AISI 4340 steel used to facilitate loading of bombs onto the underside of military aircraft could not sustain required loads during routine proof load testing. Three failed lugs underwent visual examination, chemical analysis, metallography, hardness testing, scanning electron microscopy, and energy-dispersive x-ray spectroscopy. It was determined that the failures were due to forging defects. Both forging laps and seams acted as stress concentrators when the lugs were loaded during proof testing.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"20 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":"125159379","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.c0006402
New aircraft wing panels extruded from 7075-T6 aluminum exhibited an unusual pattern of circular black interrupted lines, which could not be removed by scouring or light sanding. The panels, subsequent to profiling and machining, were required to be penetrated inspected, shot peened, H2SO4 anodized, and coated with MIL-C-27725 integral fuel tank coating on the rib side. Scanning electron microscopy and microprobe analysis (both conventional energy-dispersive and Auger analyzers) showed that the anodic coating was applied over an improperly cleaned and contaminated surface. The expanding corrosion product had cracked and, in some places, had flaked away the anodized coating. The corrodent had penetrated the base aluminum in the form of subsurface intergranular attack to a depth of 0.035 mm (0.0014 in.). It was recommended that a vapor degreaser be used during cleaning prior to anodizing. A hot inhibited alkaline cleaner was also recommended during cleaning prior to anodizing. The panels should be dichromate sealed after anodizing. The use of deionized water was also recommended during the dichromate sealing operation. In addition, the use of an epoxy primer prior to shipment of the panels was endorsed. Most importantly, surveillance of the anodizing process itself was emphasized.
由7075-T6铝材挤压而成的新飞机机翼面板显示出一种不寻常的圆形黑色断线图案，这种图案无法通过冲刷或光砂去除。在成型和加工之后，面板需要进行穿透检查、喷丸处理、H2SO4阳极氧化，并在肋侧涂上MIL-C-27725整体油箱涂层。扫描电子显微镜和微探针分析(包括传统的能量色散和俄歇分析仪)表明，阳极涂层被应用在不正确清洁和污染的表面上。膨胀的腐蚀产物已经破裂，在某些地方，阳极氧化涂层已经剥落。腐蚀以表面下晶间侵蚀的形式渗入基体铝，深度为0.035 mm (0.0014 in.)。建议在阳极氧化前的清洁过程中使用蒸汽脱脂剂。在阳极氧化之前的清洁过程中，也建议使用热抑制碱性清洁剂。阳极氧化后的面板应采用重铬酸盐密封。在重铬酸盐密封过程中也建议使用去离子水。此外，批准在面板装运前使用环氧底漆。最重要的是，强调了对阳极氧化过程本身的监测。
{"title":"Corrosion of Aluminum Alloy 7075-T6 Wing Panel","authors":"","doi":"10.31399/asm.fach.aero.c0006402","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0006402","url":null,"abstract":"\u0000 New aircraft wing panels extruded from 7075-T6 aluminum exhibited an unusual pattern of circular black interrupted lines, which could not be removed by scouring or light sanding. The panels, subsequent to profiling and machining, were required to be penetrated inspected, shot peened, H2SO4 anodized, and coated with MIL-C-27725 integral fuel tank coating on the rib side. Scanning electron microscopy and microprobe analysis (both conventional energy-dispersive and Auger analyzers) showed that the anodic coating was applied over an improperly cleaned and contaminated surface. The expanding corrosion product had cracked and, in some places, had flaked away the anodized coating. The corrodent had penetrated the base aluminum in the form of subsurface intergranular attack to a depth of 0.035 mm (0.0014 in.). It was recommended that a vapor degreaser be used during cleaning prior to anodizing. A hot inhibited alkaline cleaner was also recommended during cleaning prior to anodizing. The panels should be dichromate sealed after anodizing. The use of deionized water was also recommended during the dichromate sealing operation. In addition, the use of an epoxy primer prior to shipment of the panels was endorsed. Most importantly, surveillance of the anodizing process itself was emphasized.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"27 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":"125332349","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.c0048257
Evidence of destructive pitting on the gear teeth (AMS 6263 steel) in the area of the pitchline was exhibited by an idler gear for the generator drive of an aircraft engine following test-stand engine testing. The case hardness was investigated to be lower than specified and it was suggested that it had resulted from surface defects. A decarburized surface layer and subsurface oxidation in the vicinity of pitting were revealed by metallographic examination of the 2% nital etched gear tooth sample. It was concluded that pitting had resulted as a combination of both the defects. The causes for the defects were reported based on previous investigation of heat treatment facilities. Oxide layer was caused by inadequate purging of air before carburization while decarburization was attributed to defects in the copper plating applied to the gear for its protection during austenitizing in an exothermic atmosphere. It was recommended that steps be taken during heat treatment to ensure neither of the two occurred.
{"title":"Failure of a Carburized Steel Generator-Drive Idler Gear by Pitting Due to Decarburization and Subsurface Oxidation","authors":"","doi":"10.31399/asm.fach.aero.c0048257","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0048257","url":null,"abstract":"\u0000 Evidence of destructive pitting on the gear teeth (AMS 6263 steel) in the area of the pitchline was exhibited by an idler gear for the generator drive of an aircraft engine following test-stand engine testing. The case hardness was investigated to be lower than specified and it was suggested that it had resulted from surface defects. A decarburized surface layer and subsurface oxidation in the vicinity of pitting were revealed by metallographic examination of the 2% nital etched gear tooth sample. It was concluded that pitting had resulted as a combination of both the defects. The causes for the defects were reported based on previous investigation of heat treatment facilities. Oxide layer was caused by inadequate purging of air before carburization while decarburization was attributed to defects in the copper plating applied to the gear for its protection during austenitizing in an exothermic atmosphere. It was recommended that steps be taken during heat treatment to ensure neither of the two occurred.","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":"126745093","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.c0006387
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.
{"title":"Failure by Stress-Corrosion Cracking of an Ejection Seat Swivel","authors":"","doi":"10.31399/asm.fach.aero.c0006387","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0006387","url":null,"abstract":"\u0000 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.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"90 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":"115344498","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.c0048630
Two clevis-head self-retaining bolts used in the throttle-control linkage of a naval aircraft failed on the aircraft assembly line. Specifications required the bolts to be heat treated to a hardness of 39 to 45 HRC, followed by cleaning, cadmium electroplating, and baking to minimize hydrogen embrittlement. The bolts broke at the junction of the head and shank. The nuts were, theoretically, installed fingertight. The failure was attributed to hydrogen embrittlement that had not been satisfactorily alleviated by subsequent baking. The presence of burrs on the threads prevented assembly to finger-tightness, and the consequent wrench torquing caused the actual fractures. The very small radius of the fillet between the bolt head and the shank undoubtedly accentuated the embrittling effect of the hydrogen. To prevent reoccurrence, the cleaning and cadmium-plating procedures were stipulated to be low-hydrogen in nature, and an adequate post plating baking treatment at 205 deg C (400 deg F), in conformity with ASTM B 242, was specified. A minimum radius for the head-to-shank fillet was specified at 0.25 mm (0.010 in.). All threads were required to be free of burrs. A 10-day sustained-load test was specified for a sample quantity of bolts from each lot.
一架海军飞机的节流阀控制连接装置中使用的两个槽头自固定螺栓在飞机装配线上失效。规格要求将螺栓热处理至39至45 HRC的硬度，然后进行清洗、镀镉和烘烤，以尽量减少氢脆。螺栓在头和柄的连接处断了。从理论上讲，这些螺母安装得很紧。失败的原因是氢脆没有通过随后的烘烤得到令人满意的缓解。螺纹上毛刺的存在使装配无法达到指紧性，而随之而来的扳手扭矩导致了实际的断裂。螺栓头和螺栓柄之间的圆角半径非常小，无疑加重了氢的脆化作用。为了防止再次发生，清洗和镀镉的程序被规定为低氢性质，并在205摄氏度(400华氏度)下进行适当的镀后烘烤处理，符合ASTM B 242的规定。头柄圆角的最小半径规定为0.25毫米(0.010英寸)。所有的螺纹都要求无毛刺。对每批螺栓的样品数量进行为期10天的持续载荷试验。
{"title":"Hydrogen Embrittlement of Cadmium-Plated Alloy Steel Self-Retaining Bolts in a Throttle-Control Linkage","authors":"","doi":"10.31399/asm.fach.aero.c0048630","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0048630","url":null,"abstract":"\u0000 Two clevis-head self-retaining bolts used in the throttle-control linkage of a naval aircraft failed on the aircraft assembly line. Specifications required the bolts to be heat treated to a hardness of 39 to 45 HRC, followed by cleaning, cadmium electroplating, and baking to minimize hydrogen embrittlement. The bolts broke at the junction of the head and shank. The nuts were, theoretically, installed fingertight. The failure was attributed to hydrogen embrittlement that had not been satisfactorily alleviated by subsequent baking. The presence of burrs on the threads prevented assembly to finger-tightness, and the consequent wrench torquing caused the actual fractures. The very small radius of the fillet between the bolt head and the shank undoubtedly accentuated the embrittling effect of the hydrogen. To prevent reoccurrence, the cleaning and cadmium-plating procedures were stipulated to be low-hydrogen in nature, and an adequate post plating baking treatment at 205 deg C (400 deg F), in conformity with ASTM B 242, was specified. A minimum radius for the head-to-shank fillet was specified at 0.25 mm (0.010 in.). All threads were required to be free of burrs. A 10-day sustained-load test was specified for a sample quantity of bolts from each lot.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"72 2 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":"116393462","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.c9001746
Jivan B. Shah
Brittle intergranular fracture, typical of a hydrogen-induced delayed failure, caused the failure of an AISI 4340 Cr-Mo-Ni landing gear beam. Corrosion resulting from protective coating damage released nascent hydrogen, which diffused into the steel under the influence of sustained tensile stresses. A second factor was a cluster of non-metallic inclusions which had ‘tributary’ cracks starting from them. Also, eyebolts broke when used to lift a light aircraft (about 7000 lb.). The bolt failure was a brittle intergranular fracture, very likely due to a hydrogen-induced delayed failure mechanism. As for the factors involved, cadmium plating, acid pickling, and steelmaking processes introduce hydrogen on part surfaces. As a second contributing factor, both bolts were 10 Rc points higher in hardness than specified (25 Rc), lessening ductility and notch toughness. A third factor was inadequate procedure, which resulted in bending moments being applied to the bolt threads.
{"title":"Hydrogen Embrittlement of Aircraft Components","authors":"Jivan B. Shah","doi":"10.31399/asm.fach.aero.c9001746","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001746","url":null,"abstract":"\u0000 Brittle intergranular fracture, typical of a hydrogen-induced delayed failure, caused the failure of an AISI 4340 Cr-Mo-Ni landing gear beam. Corrosion resulting from protective coating damage released nascent hydrogen, which diffused into the steel under the influence of sustained tensile stresses. A second factor was a cluster of non-metallic inclusions which had ‘tributary’ cracks starting from them. Also, eyebolts broke when used to lift a light aircraft (about 7000 lb.). The bolt failure was a brittle intergranular fracture, very likely due to a hydrogen-induced delayed failure mechanism. As for the factors involved, cadmium plating, acid pickling, and steelmaking processes introduce hydrogen on part surfaces. As a second contributing factor, both bolts were 10 Rc points higher in hardness than specified (25 Rc), lessening ductility and notch toughness. A third factor was inadequate procedure, which resulted in bending moments being applied to the bolt threads.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"17 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":"114472013","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.c9001552
Jivan B. Shah
This report covers case histories of failures in fixed-wing light aeroplane and helicopter components. A crankshaft of AISI 4340 Ni-Cr-Mo alloy steel, heat treated and nitrided all over, failed in bending fatigue. The nitrided layer was ground too rapidly causing excessive heat generation which induced grinding cracks and grinding burn. Tensional stresses resulting from grinding developed in a thin surface layer. On another crankshaft, chromium plating introduced undesirable residual tensile stresses. Such plating is an unsatisfactory finish for crankshafts of aircraft engines. Aircraft engine manufacturers and aeronautical standards require magnetic particle inspection to detect grinding cracks after reconditioning. Renitriding after any grinding is needed also, regardless of the amount of undersize as it introduces beneficial residual compressive stresses.
{"title":"Failures in Airplane Engine Crankshafts","authors":"Jivan B. Shah","doi":"10.31399/asm.fach.aero.c9001552","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c9001552","url":null,"abstract":"\u0000 This report covers case histories of failures in fixed-wing light aeroplane and helicopter components. A crankshaft of AISI 4340 Ni-Cr-Mo alloy steel, heat treated and nitrided all over, failed in bending fatigue. The nitrided layer was ground too rapidly causing excessive heat generation which induced grinding cracks and grinding burn. Tensional stresses resulting from grinding developed in a thin surface layer. On another crankshaft, chromium plating introduced undesirable residual tensile stresses. Such plating is an unsatisfactory finish for crankshafts of aircraft engines. Aircraft engine manufacturers and aeronautical standards require magnetic particle inspection to detect grinding cracks after reconditioning. Renitriding after any grinding is needed also, regardless of the amount of undersize as it introduces beneficial residual compressive stresses.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"30 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":"116354747","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.c0092142
During a routine inspection on an aircraft assembly line, an airframe attachment bolt was found to be broken. The bolt was one of 12 that attach the lower outboard longeron to the wing carry-through structure. Failure occurred on the right-hand forward bolt in this longeron splice attachment. The bolt was fabricated from PH13-8Mo stainless steel heat treated to have an ultimate tensile strength of 1517 to 1655 MPa (220 to 240 ksi). A water-soluble coolant was used in drilling the bolt hole where this fastener was inserted. Investigation (visual inspection, 265 SEM images, hardness testing, auger emission spectroscopy and secondary imaging spectroscopy, tensile testing, and chemical analysis) supported the conclusion that failure of the attachment bolt was caused by stress corrosion. The source of the corrosive media was the water-soluble coolant used in boring the bolt holes. Recommendations included inspecting for corrosion all the bolts that were installed using the water-soluble coolant at the spliced joint areas, rinsing all machined bolt holes with a noncorrosive agent, and installing new PH13-8Mo stainless steel bolts with a polysulfide wet sealant.
{"title":"Bending of an Aircraft-Wing Slat Track","authors":"","doi":"10.31399/asm.fach.aero.c0092142","DOIUrl":"https://doi.org/10.31399/asm.fach.aero.c0092142","url":null,"abstract":"\u0000 During a routine inspection on an aircraft assembly line, an airframe attachment bolt was found to be broken. The bolt was one of 12 that attach the lower outboard longeron to the wing carry-through structure. Failure occurred on the right-hand forward bolt in this longeron splice attachment. The bolt was fabricated from PH13-8Mo stainless steel heat treated to have an ultimate tensile strength of 1517 to 1655 MPa (220 to 240 ksi). A water-soluble coolant was used in drilling the bolt hole where this fastener was inserted. Investigation (visual inspection, 265 SEM images, hardness testing, auger emission spectroscopy and secondary imaging spectroscopy, tensile testing, and chemical analysis) supported the conclusion that failure of the attachment bolt was caused by stress corrosion. The source of the corrosive media was the water-soluble coolant used in boring the bolt holes. Recommendations included inspecting for corrosion all the bolts that were installed using the water-soluble coolant at the spliced joint areas, rinsing all machined bolt holes with a noncorrosive agent, and installing new PH13-8Mo stainless steel bolts with a polysulfide wet sealant.","PeriodicalId":326464,"journal":{"name":"ASM Failure Analysis Case Histories: Air and Spacecraft","volume":"20 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":"114806125","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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