Folkston, GA, USA
N3096H
Ayres S2R-T34
The pilot observed black smoke and white sparks from the engine exhaust while climbing between 80 to 100 feet on initial takeoff climb followed by a sudden loss of engine power. The pilot landed straight ahead to none suitable terrain. The landing gear separated and the bottom of the fuselage received structural damage. Examination of the airplane reveled separated turbine blades in the exhaust section of the airplane. The engine was disassembled. The compressor turbine shroud exhibited battering and gouges. The compressor turbine disc blade airfoils were fractured from the root to about three quarters span and exhibited battering and gouging. The disc upstream side blade platforms exhibited circumferential rubbing. The compressor turbine disc was forwarded to the Pratt & Whitney Materials Laboratory for examination under the supervision of Transport Safety Board Canada . The examination revealed the compressor turbine blades failed in overload due to high temperature exposure.
On April 5, 2006, at 1821 eastern daylight time, an Ayers S2R-T34, N3096, registered to and operated by Souther Crop Spraying Inc., as a 14 CFR Part 137 aerial application flight, experienced a total loss of engine power during climb after takeoff from Davis Field, Folkston, Georgia. Visual meteorological conditions prevailed and no flight plan was filed. The airplane received substantial damage. The airline transport pilot reported minor injuries. The flight was originating at the time of the accident. The pilot stated prior to departing from runway 18 all engine instruments were normal. After rotation the pilot observed black smoke and white sparks from the engine exhaust while climbing between 80 to 100 feet. There was a sudden loss of engine power and the pilot landed straight ahead to non suitable terrain. The landing gear separated and the bottom of the fuselage received structural damage. Initial examination of the airplane by the FAA reveled separated turbine blades in the exhaust section of the airplane. The airplane was recovered and the engine was removed from the airframe on April 13, 2006. The engine was stripped of all airframe components in the presence of the FAA. All engine related controls and accessories were in place and intact. The engine was transported to a Pratt & Whitney Service Center in Atlanta, Georgia ,for further examination in the presence of the NTSB on April 20, 2006. All positional references during the engine examination are in relation to view from aft looking forward. Upstream and downstream references are in relation to gas path flow from the compressor inlet to exhaust. The reduction gearbox housing was intact and the oil scavenge transfer tube boss was fractured. The exhaust duct housing displayed compressional deformation. The gas generator case displayed compressional deformation around the upper circumference between the engine mount collar and the fuel nozzle ports. The accessory gearbox housing was intact. The forward beta linkage arm was not received with the engine for the examination, and all remaining linkage was continuous and intact The compressor discharge air (P3) was intact and continuous from the gas generator case fitting to the fuel control unit fitting. The power turbine control (Py) was intact and continuous from the fuel control unit fitting to the propeller governor fitting. The reduction gearbox chip detector displayed ferrous debris and two small ball bearings about five sixteenths of an inch in diameter. The oil filter was removed and was free of contaminants. The gas generator rotor moved freely when turned by hand and was continuous with all accessory drives. The propeller shaft was rotated by hand and the power turbine was seized. The compressor first stage compressor exhibited no distress and the compressor was not disassembled. The combustion chamber liner exhibited no indications of operational distress. The large exit duct exhibited no indication of operational distress and the flame pattern was normal. The small exit duct exhibited no indication of operational distress; however erosion and cracking was present at the 12:00 position. The small exit duct was forward to the Pratt & Whitney Materials Laboratory for further examination under the supervision of Transportation Safety Board (TSB) Canada. The compressor turbine guide vane ring airfoil trailing edges exhibited battering, gouges, tearing, and burning. The vane leading edges exhibited no burning and minor coating loss. The anti rotation lugs were fractured. The compressor turbine guide vane ring and small exit duct assembly were forwarded to the Pratt & Whitney Materials Laboratory for further examination under the supervision of TSB Canada. The compressor turbine shroud exhibited battering, and gouges. The compressor turbine disc blades (CT) were numbered 1 through 58 in a clockwise direction in relation to the hub master spline. The blade airfoils were fractured from the root to about three quarters span and exhibited battering and gouging. The disc upstream side blade platforms exhibited circumferential rubbing between blade position 3 and 18, and the downstream side blade platforms exhibited circumferential rubbing. The compressor turbine disc was forwarded to the Pratt & Whitney Materials Laboratory for further examination under the supervision of TSB Canada. The interstate turbine temperature probes were fractured, and the power turbine housing was intact. The Pratt and Whitney Materials Investigation-Laboratory Report stated, "The airfoil of the CT blades failed by overload. The microstructure showed evidence of high temperature exposure. The microstructure also showed evidence of rafting suggesting a creep mechanism. However, there was no evidence of surface necking at the surface of the airfoils adjacent to the fractures that would have confirmed that the blades fractured by a creep fracture. Metallographic examination of a longitudinal section taken at approximately one quarter of an inch from the trailing edge showed a significant rafting of the gamma prime precipitates above the mid-span of the airfoil. No evidence of micro-void or decohesion was observed. The root cause of blades distress could not be determined with certainty because the first blade to be released could not be identified." The power turbine guide vane ring and interstage baffle vane airfoils were battered and gouged. The interstage baffle upstream side exhibited circumferential rubbing. The downstream side baffle face was rubbed and deformed. The power turbine shroud was battered, gouged, and exhibited circumferential rubbing. The power turbine disc was displaced to about the 12:00 position against the power turbine shroud. About two thirds of the blade airfoils were fractured. Several blade tips were recovered within the exhaust duct. The blade airfoil leading edges exhibited battering and gouging. The blade tip shrouds exhibited circumferential rubbing. The power turbine shaft and shaft housing was not disassembled. The reduction gearbox propeller shaft rotated with resistance, and was not coupled with the seized power turbine. The reduction gearbox was separated at flange "A" for further examination. The flexible coupling to the first stage splined adapter retaining rings, and the flexible coupling to the second stage sun gear retaining ring were separated from their seats. The rear land of the flexible coupling retaining ring seat exhibited circumferential rubbing. The first and second stage gearing exhibited no indications of operational distress. The accessory gearbox rotated freely by hand and was continuous with the gas generator rotor. The accessory gearbox was not disassembled. The engine controls and accessories were not removed for further examination. The components retained for further examination by the Pratt and Whitney Materials Laboratory were released to Atlanta Air recovery on August 16, 2006.
The total loss of engine power due to the failure of the compressor turbine blades due to high temperature exposure.
Source: NTSB Aviation Accident Database
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