Yuba City, CA, USA
N7503L
SCHWEIZER AIRCRAFT CORP G-164B
The pilot reported that, during the initial climb, he heard a "loud pop" and observed fire coming from both exhaust stacks before the engine lost power. He subsequently made an off-airport landing to an adjacent road. During the landing, the airplane slid into a ditch, the landing gear sheared off, and the airplane nosed over. A postaccident engine examination revealed that the blade retaining ring of the 1st-stage compressor blades had fractured into several pieces; about 30 percent of the ring was not recovered. A metallurgical examination of the recovered sections of the retaining ring revealed that all but two of the fractures were due to ductile overload. An examination of the nonoverload fractures revealed fatigue originating from the rivet hole areas, which was indicative of high-cycle fatigue. The power section had been overhauled 747 hours before the accident. According to the overhaul manual current at the time of the overhaul, the blade retaining ring was an on-condition inspection item and not required to be replaced at overhaul. The fatigue fracture of the 1st-stage compressor blade retaining ring likely affected the engine's ability to produce power.
HISTORY OF FLIGHTOn July 9, 2011, about 1430 Pacific daylight time, a Schweizer G-164B, N7503L, encountered a ditch during an off airport forced landing following a loss of engine power near Yuba City, California. Onstott Dusters, Inc., was operating the airplane under the provisions of 14 Code of Federal Regulations (CFR) Part 137. The commercial pilot was not injured; the airplane sustained substantial damage from impact forces. The local aerial application flight was departing from a private dirt strip. Visual meteorological conditions prevailed, and no flight plan had been filed. The pilot stated that the airplane reached about 35 feet above ground level after takeoff when he heard a loud pop. He observed fire coming from both exhaust stacks, and the engine lost power. He landed on an adjacent road, but the airplane slid into a ditch alongside it. The landing gear sheared off, and the airplane nosed over. AIRCRAFT INFORMATIONThe airplane was a Schweizer G164B Ag Cat, serial number 826B. The operator reported that the airplane had a total airframe time of 10,282 hours at the last annual inspection on April 19, 2011. The engine was a Pratt & Whitney Canada (PWC) PT6A-34, serial number PCE 56022. Total time recorded on the engine at the last 100-hour annual inspection was 8,878 hours. The engine had 747.7 hours since overhaul on March 5, 2009. AIRPORT INFORMATIONThe airplane was a Schweizer G164B Ag Cat, serial number 826B. The operator reported that the airplane had a total airframe time of 10,282 hours at the last annual inspection on April 19, 2011. The engine was a Pratt & Whitney Canada (PWC) PT6A-34, serial number PCE 56022. Total time recorded on the engine at the last 100-hour annual inspection was 8,878 hours. The engine had 747.7 hours since overhaul on March 5, 2009. ADDITIONAL INFORMATIONPWC updated section 72-30-05, page 107, paragraph B(4) of the engine Overhaul Manual on August 8, 2013, to indicate that the blade retaining ring should be discarded at disassembly. TESTS AND RESEARCHPWC examined the engine under the supervision of a Federal Aviation Administration (FAA) inspector at the PWC Service Investigation Facilities in Bridgeport, West Virginia on January 11, 2012, and the PWC materials laboratory completed a follow up metallurgical exam. A report of the findings is part of the public docket for this accident. Pertinent excerpts follow. No debris was observed on the magnetic poles of the reduction gearbox chip detector. The oil filter was clean, and the residual oil in the filter and housing was clean. The fuel filter was clean, and residual fuel in the filter and housing was clean. Compressor Section The 1st stage compressor blades had damage on the leading and trailing edge of the airfoils. The blade retaining ring had fractured into several pieces; approximately 30 percent of the ring was not recovered. The rear hub (1st stage rotor) was in good condition; four of the six required blade retaining ring's retention rivets were in place, but the flared end had fractured against the hub. The PWC metallurgist's examination of the recovered sections of the retaining ring revealed that all but two of the fractures were due to ductile overload. A Scanning Electron Microscope (SEM) examination of one of the non-overload fractures exhibited fatigue that originated from the rivet hole area. The fatigue cracks were indicative of high cycle fatigue (HCF) from the rivet hole. The area on the opposite side of the rivet hole fractured in shear. A SEM examination of the second fatigue crack location also indicated HCF. No material anomalies were found at the fatigue initiation site area, and the chemical composition of the ring met manufacturing drawing requirements. Per the current Overhaul Manual (P/N 3021243 rev 31, dated December 15, 2006) at the time of the overhaul, the blade retaining ring was an on-condition at inspection item, and not required to be replaced at overhaul. The 2nd and 3rd stage compressor blades exhibited leading edge damage, the tips were battered, and there was trailing edge deformation. The compressor disk was in good condition. Twelve 1st stage compressor stator airfoils fractured at the housing interface, and were liberated from the housing. The remaining airfoils were bent in the direction of compressor rotation, and in an upstream direction away from the 2nd stage rotor. Analysis by the PWC metallurgist revealed that all of the stator airfoils fractured in a ductile manner under shear impact. The chemical composition of the stators met drawing requirements. The hardness of two of the stators was slightly lower than manufacturer's drawing requirements; the other stators met drawing requirements. The 2nd stage compressor stator exhibited nine fractured and liberated airfoils. The remaining airfoils were all bent in the direction of the compressor rotation, and were also bent in a downstream direction away from the 2nd stage rotor. There was damage to both the leading and trailing edges of the airfoils. The 3rd stage compressor stator exhibited one fractured and liberated airfoil. The remaining airfoils exhibited varying degrees of damage on both sides of the airfoil, bent in different directions/angles, and had tip rub with its respective spacer. All three spacers rubbed their respective mating vane tips. The centrifugal impeller had leading edge damage with heavy rub indications along the entire airfoil profile. The centrifugal impeller shroud had a rub mark with material pick up from the impeller. The number one bearing was wet and rotated normally; there was a light rub on the air seals. The number two bearing was wet and rotated normally; the rotor air seal was discolored with displaced material from contact with the stator air seal. Combustion Section The combustion chamber liner was bent, distorted, and twisted; the interior was covered with oil and black carbon. Two areas of the dome were bent. Turbine Section Numerous components of the turbine section were dented and damaged. There were rub marks on various adjacent components. The compressor turbine shroud exhibited a heavy rub mark with both displaced material and metal splatter build up. The power turbine center bore area exhibited dents and distortions. The balancing rim rubbed the PT vane baffle and the anti-rotation lugs were bent. The blade serrations rubbed the baffle. The blades were loose and fractured at varying heights. Metallurgy examination of two randomly selected blades indicated that they exhibited tensile overload fractures. No material or overtemperature anomalies were observed; the chemical composition of the blades met drawing requirements. The number three and four bearings of the power turbine shaft and shaft housing were wet and rotated normally, but were stained. The shaft was unremarkable except that it was stained. The number three bearing air seal flange and the number four retention flange bulged forward. Reduction Gearbox The front and rear gearboxes were not disassembled. Both gearboxes rotated freely after the power turbine shaft housing was removed. No visual anomalies were noted. Accessory Gearbox The gearbox was not disassembled; it rotated freely after it was removed from the engine. No visual anomalies were noted.
High-cycle fatigue on the 1st-stage compressor blade retaining ring, which resulted in the ring’s fracture and a subsequent loss of engine power.
Source: NTSB Aviation Accident Database
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