Santa Rosa, CA, USA
N2502C
Cessna 170B
The engine lost power and the airplane nosed over during a forced landing in a field. Post accident examination revealed that the number 2 cylinder separated from the engine. The engine had accumulated 900 hours since its last major overhaul 13 years prior to the accident; the cylinder had accumulated 1,212 hours since replacement in 1986. The hours on the cylinder prior to that time could not be determined. A metallurgical examination found that the barrel failed due to a fatigue crack. The barrel fractured circumferentially between the 11th and 13th fin from the head. The fracture initiated approximately 1.8 inches below the head and started from the inner diameter below a plated region. The initiation area was a small, flat, thumbnail-shaped region approximately 0.5-inch wide through the wall thickness, with signatures indicative of fatigue. Multiple crack initiation sites were observed in the origin area. From the initiation point, the fracture continued propagating in the circumferential direction with coarse arrest marks over approximately 45 percent of the barrel circumference. No material, manufacture, or plating defects were found. The metallurgical examination could not determine why the cracking occurred at the given location; however, it occurred below the chrome plating. From this, it is possible that the barrel exhausted the available fatigue life at that location, which was lowered by the presence of plating.
On January 17, 2004, at 1525 Pacific standard time, a Cessna 170B, N2502C, force landed in a field 15 miles south of Santa Rosa Airport (STS), Santa Rosa, California, after a loss of engine power. The aircraft nosed over in soft ground during the landing rollout. The pilot, who was also the registered owner, was operating the airplane under the provisions of 14 CFR Part 91. The private pilot, the sole occupant, was not injured; the airplane sustained substantial damage. The flight departed Santa Rosa about 1505 for the local flight. Visual meteorological conditions prevailed, and no flight plan had been filed. A Federal Aviation Administration (FAA) inspector examined the airplane. The number 2 cylinder separated from the engine, but remained inside the engine cowling. The engine had accumulated 900 hours since its last major overhaul. The National Transportation Safety Board investigator-in-charge (IIC) reviewed maintenance records for the airplane. On September 4, 1986, the number 2 cylinder was replaced. On April 22, 2003, the number 2 cylinder was removed and resealed. The cylinder had accumulated 1,212 hours since replacement in 1986. The hours on the cylinder prior to that time could not be determined. On May 20, 1991, the engine was overhauled. The IIC sent the outboard portion of the number 2 cylinder to the Safety Board Materials Laboratory for examination. According to the metallurgist, the barrel fractured circumferentially between the 11th and 13th fin from the head. Examination of the fracture surface showed that the fracture initiated approximately 1.8 inches below the head and started from the inner diameter below a plated region. The origin was on the intake valve side of the cylinder, at approximately the 9-o'clock position looking outward when the rocker arms are at the 6 o'clock position. The initiation area was a small, flat, thumbnail-shaped region approximately 0.5-inch wide through the wall thickness, indicative of fatigue. The metallurgist noted multiple initiation sites in the origin area. From this region the fracture continued propagating in the circumferential direction with coarse arrest marks. The total fatigue region consumed approximately 45 percent of the barrel circumference or approximately 5.5 inches. The remainder of the barrel failed in overstress. Energy dispersive spectroscopy (EDS) analysis of the base metal showed a composition consistent with 4140 steel, a typical barrel material, and the plating was composed of chromium. The metallurgist made an as-polished cross-sectional metallographic mount through the origin area and no material defects were observed. At the inner diameter surface near the origin, he observed a secondary crack propagating adjacent to the main fracture. The etched microstructure was consistent with through-hardened quenched and tempered steel. The region on the surface unaffected by etching is the chromium plating, which contained multiple craze cracks. Sections of the plating adjacent to the fracture origin were staggered. The plating thickness was approximately 0.012 inch, whereas, the base metal thickness in the thinnest region in the fin root was approximately 0.073 inch. The metallurgist stripped the barrel inner diameter surface of the chrome plating by immersing the part in a 50 percent solution of hydrochloric acid to determine the surface conditions adjacent to the fatigue origins. He observed no evidence of abusive machining marks, scratches, or corrosion pits in the origin areas, though some deeper machining marks were observed to have been removed from the origins. He observed additional areas of secondary cracking adjacent to the fracture surface. The barrel had an average micro-hardness of 328 HK (32 HRC) in the base metal adjacent to the origin, consistent with typical barrel hardness. The chrome plating had an average hardness of 813 HK (63 HRC).
the failure of the number 2 cylinder barrel due to fatigue. The lack of suitable landing terrain was a factor in the accident.
Source: NTSB Aviation Accident Database
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