VALLEY SPRINGS, CA, USA
N38908
Piper PA-32R-300
The airplane had a catastrophic engine failure while the pilot was in cruise flight at 4,500 feet. The pilot declared an emergency and began to perform the emergency checklist. The pilot located a flat field and struck a barbed wire fence during the landing roll, which damaged the wings and collapsed the nose gear. The engine was sent back to the Textron Lycoming factory for materials examination. The No. 6 connecting rod initially fractured through the cap with the fracture most likely initiating from an area of galling on the inner diameter of the cap. The fracture was propagated by a fatigue mechanism until it failed with catastrophic overload. Following the metallurgical examination on the No. 6 connecting rod and associated components, additional disassembly of the engine was accomplished that did not conform to investigative protocols. No measurements were taken of the case through bolt torque or the remaining rod end cap bolt torque. No determination was made whether the bearing inserts were of the correct size or assembled properly in engine build-up. Review of the engine maintenance records disclosed that it was remanufactured by the Lycoming factory in 1991, and had operated 869 hours since rebuild and installation in the airframe. No removal and replacement of cylinders was noted in the records.
On May 28, 1999, at 1623 hours Pacific daylight time, a Piper PA-32R-300, N38908, collided with a fence during a forced landing following a catastrophic engine failure in cruise flight near Valley Springs, California. The commercial pilot/owner, the sole occupant, was not injured. The aircraft sustained substantial damage during the accident sequence. The business flight was operated by the pilot/owner under 14 CFR Part 91. No flight plan was filed for the cross-country flight that had departed Atwater Castle airport at 1600 and was destined for Cameron Park, California. Visual meteorological conditions prevailed at the time of the accident. The pilot stated that be did a preflight and that everything appeared to be normal. Nine quarts of oil and approximately 84 gallons of fuel were onboard. He said he took off and climbed to 4,500 feet, trimmed the airplane, and set the engine for 2,400 rpm and 24.6 inches of manifold pressure at a 17 gallons-per-hour fuel consumption rate. He said that the engine was running smooth. About 1615, the engine began running very rough, and then lost power. The pilot said he declared an emergency with Oakland Center, and began to perform an emergency checklist. He said he heard a "pop" sound and oil began splattering all over the windshield and the engine abruptly stopped turning. He estimated the time from the rough engine to the engine seizure about 2 minutes. He said he located a large relatively flat field and told Oakland Center that he was going to land. He said he proceeded to turn off all switches, fuel flow, and ignition. A circling approach was made to the field using full flaps, and the airplane struck a barbwire fence during the landing roll, which damaged the wings and collapsed the nose gear. The Calveras County Sheriff's department responded to the scene and took the pilot's statement. According to the police report, as the pilot began his approach on the north side of Hogan Dam Road, the landing roll was longer than expected, and the airplane continued to roll, which caused the airplane to go through a wire fence. The airplane knocked down about 40-feet of fence and fence posts. One post hit each wing, which caused about 6-8 feet of damage to the leading edge of each wing. Federal Aviation Administration (FAA) inspectors who arrived at the scene observed that the engine case was damaged and that there was a connecting rod protruding from the case. Arrangements were made to have the airplane removed from the field and taken to a secure location for further examination. On June 4, 1999, at the request of Safety Board investigators, FAA inspectors and a technical representative from Textron Lycoming examined the engine at the retrieval site. The inspection revealed that the Nos. 4 and 6 cylinders were cracked and displaced outward. Additionally, the No. 6 connecting rod was partially protruding from the top of the engine case. Fragments of the connecting rod and one of the two-rod bolts were found resting on the baffle between the Nos. 4 and 6 cylinder. The top spark plugs were removed and examined. According to the Textron Lycoming representative, the spark plugs remained undamaged and displayed coloration consistent with normal operation. Additionally, the cylinder combustion chambers were examined using a lighted bore scope. The valves were found to be intact and undamaged. Approximately 6 quarts of oil was drained from the sump drain, and the oil suction screen was removed for examination. The screen was found to be full of loose pieces of bearing material, aluminum casting, and steel materials. The Lycoming representative removed portions of the crankcase that was cracked and displaced to expose the internal components of the engine. The Nos. 5 and 6 cylinder assemblies were removed to facilitate the examination. The No. 6 connecting rod was detached at the crankshaft journal. The examination revealed that the cylinder barrel skirt sustained damage along the rotational plane of the connecting rod. The crankshaft's No. 6 rod journal remained intact and there were no signatures consistent with heat distress observed at the journal. According to the maintenance records, the airplane's engine, a Lycoming IO-540-KIG5D, had 869 hours since factory remanufacture in 1991. Following the rebuild, the engine was installed in the airframe in November of that year. No major repair or removal and replacement of cylinders were noted in the records. The engine was then sent back to the Textron Lycoming Materials Laboratory for further examination under the supervision of FAA inspectors. In the Textron Lycoming Materials Laboratory report, visual examination of the No. 6 cylinder revealed that the piston was jammed into the cylinder dome. Additionally, there was extensive damage to the yoke end of the connecting rod. The bolting areas on the rod side of the yoke had broken off. The cap of the connecting rod had broken into at least three pieces, which were subsequently submitted to the materials laboratory. Examination of the pieces of the connecting rod yoke area revealed no thermal discoloration and no evidence of extruded connecting rod bearings. Fractographic examination of the undamaged areas on the pieces revealed an overload mechanism on all fractures with one exception. A fracture surface on one of the cap pieces exhibited fatigue propagation as evidenced by the transgranular topography, fracture rays, and numerous beach marks. The origin area was located on the cap inner diameter in the area adjacent to the cutout for the connecting rod bolt. According to the Textron laboratory report, the origin area and cap inner diameter had severe secondary mechanical damage; however, based upon the pattern of the fracture rays the metallurgist concluded that the fracture initiated from a single origin area. Examination of the cap inner diameter revealed galling adjacent to the other bolt cutout. The Nos. 1, 2, and 5 connecting rods also revealed galling on one or both sides of the cap at this location. Connecting rod caps Nos. 3 and 4 were too damaged for evaluation of the cap inner diameter. According to the report, the microstructure of the connecting rod consisted of tempered martensite, which is typical of properly heat-treated AMS 6322. No areas of decarburization were observed. The hardness of the connecting rod was 31 HRC, which conforms to the 29-33 HRC engineering drawing requirement. The chemistry was found to conform to the AMS 6322 engineering drawing requirement. The laboratory report stated that based on the examinations, galling between the connecting rod cap and the connecting rod bearing was the most likely the source of the crack initiation. The Textron Lycoming Materials laboratory concluded that the connecting rod bolt had fractured by a ductile tensile overload mechanism. The second bolt had bent. Examination of the threads on the bolts and nuts revealed no damage and indicated that they were most likely properly torqued. The microstructure of the fractured bolt and associated nut consisted of tempered martensite, which is typical of properly heat-treated AMS 6322. The hardness of the bolt was 46 HRC, which conforms to the 46-50 HRC engineering drawing requirement. The hardness of the nut was 32 HRC, which conforms to the 28-32HRC engineering drawing requirement. The chemistry of the bolt and nut both conformed to the AMS 6322 engineering drawing requirement. Following the metallurgical examination on the No. 6 connecting rod and associated components, additional disassembly of the engine was accomplished that did not conform to investigative protocols. No measurements were taken of the case through bolt torque or the remaining rod end cap bolt torque. No determination was made whether the bearing inserts were of the correct size or assembled properly in engine build-up. The airplane was released to the insurance company, representing the registered owner, on November 19, 1999.
The fatigue failure and separation of the No. 6 connecting rod end cap, which led to a catastrophic failure of the engine. The fatigue crack initiation was due to galling on the rod end cap, which was most likely caused by an undetermined factory manufacturing process error during the engine rebuild.
Source: NTSB Aviation Accident Database
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