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On January 10, 2006, at 0845 Hawaiian standard time, a Eurocopter AS350BA, N3607P, experienced a loss of engine power and autorotated into a dense tropical forest on a steep slope near Hana, Maui, Hawaii. Sunshine Helicopters, Inc., operated the helicopter as a scenic tour flight under the provisions of 14 CFR Part 135. The four passengers received minor injuries and the airline transport pilot was seriously injured. The helicopter was destroyed. Visual meteorological conditions prevailed, and a company flight plan had been filed. The sightseeing flight duration is about an hour and returns to the departure airport. The scenic tour flight departed Kahului Airport, Maui, at 0807.
The pilot stated that following a viewing of Haleakala Crater, they proceeded to enter the Manawainui Gulch to view the waterfalls there. The pilot maneuvered the helicopter to exit the gulch and proceeded north over the Maui rainforest. At this time the helicopter vibrated, shuddered, and the low rotor rpm warning horn sounded. The pilot entered an autorotation and attempted to arrest the helicopter's forward velocity before settling into the canopy of treetops. The helicopter dropped nose first towards the forest floor, and came to rest on its right side suspended in the trees a few feet from the ground. The four passengers and the pilot were able to lower themselves to the ground and called for assistance using a cell phone.
The Eurocopter AS350BA helicopter is a six passenger, single pilot, single engine aircraft, with a traditional main rotor and tail rotor configuration. The operator reported a total airframe time of 14,007 hours at the time of the accident, and a 100-hour inspection had been completed on December 28, 2005.
The helicopter was equipped with a video recording system that could record images outside the helicopter and inside the cockpit. The video recording utilized VHS cassette tape and compact flash memory. The video recording system was operating at the time of the accident.
The engine was a Turbomeca Arriel 1B, serial number 4467. Examination of the engine logbook revealed that the engine had operated for 9,593 hours since new, and 1,764 hours since the last overhaul. The engine maintenance logbook documented that Module 3, the gas generator section, was removed, and replaced with an overhauled module on September 14, 2004. The overhaul on the replacement module had been completed by the ACRO Helipro Company on August 23, 2004. The overhaul included the replacement of the second stage turbine blades with overhauled turbine blades. Turbomeca specified the life limit of turbine blades for the Arriel 1B to be 6,000 hours. The overhauled blade set was documented to have 2,986.2 hours at the time of the turbine overhaul. The module had accumulated an additional 1,763.5 hours, and 1,712 cycles since the overhaul.
WRECKAGE & IMPACT INFORMATION
The wreckage was recovered to the Kahului Airport and examined by a National Transportation Safety Board investigator on January 13, 2006. An onboard video recording was removed from the wreckage and sent to the Safety Board Vehicle Recorders Laboratory, Washington, D.C., for further documentation and viewing. The engine was removed and crated, under the supervision of the Safety Board investigator, for shipment to the manufacturer in order to conduct a detailed teardown and examination.
TESTS & RESEARCH
On January 26, 2006, the engine was examined by representatives from Turbomeca USA, Sunshine Helicopters, ACRO Helipro, American Eurocopter, and the Federal Aviation Administration (FAA), under the supervision of the Safety Board investigator-in-charge (IIC), at the Turbomeca facility, Grand Prairie, Texas. Details of the examination are in the public docket of this report.
The engine was removed from the shipping crate and placed on an engine stand. The engine had localized sooting with no foreign debris observed. The exhaust duct exhibited a large amount of interior dents. Dents varied in size from 2 mm to 10 mm.
Upon disassembly of module 3, the gas generator high pressure turbine section, it was observed that the second stage turbine (T2) blade tips all exhibited fractures, and a single blade had separated at the blade platform flush with the disk circumference. The missing blade fir tree base was not displaced axially, the fracture surface was gray in color and granular in texture. The first stage turbine (T1) exhibited minor blade tip rubbing. All turbine blades and blade locks were present on the power (free) turbine. The blades exhibited substantial mechanical damage to all blade tips. The T2 turbine wheel and blades were removed from module 3 and sent to the Safety Board Materials Laboratory, Washington, D.C., for further metallurgical analysis.
Materials Laboratory Examination
The 2nd stage turbine disk with fractured blades was examined at the Safety Board Materials Laboratory. The disk face was labeled with part number 0292253250 and serial number ADE 2318YC. Engine records indicated that disk ADE2318YC was part of gas generator (module 3), s/n 4801TEC, and had accumulated 4,749.8 hours and 4,589 cycles since new. Module 3, 4801TEC, had accumulated 1,763.5 hours, and 1,712 cycles since its overhaul and installation in this engine, in September 2004. All blades were tight in the disk with the locks in-place and properly deformed.
All of the 2nd stage blades, except one, were fractured through the airfoil sections. The excepted blade was fractured through the "fir tree" area just below the platform. Microscopic examinations also uncovered a crack in the fir tree of another nearby blade. Each blade was marked with a sequence number 1 thru 27. However, the sequence numbers were not in order nor did they match the disk slot numbers. The fractured blade was marked 23 and the cracked blade was marked 25.
Examinations of the airfoil fractures revealed features indicative of overstress separations. Close inspection of the blade (23) fracture through the fir tree area uncovered a darker oxidized region which contained features consistent with fatigue propagation followed by overstress separation. The plane of the fracture was in the most outboard fir tree radius just below the blade platform.
The fatigue region was at the aft (trailing edge) side of the blade below the platform. Fracture marking indicated fatigue initiation at the aft trailing corner in the cutout portion of the fir tree. As viewed with the blade installed in the disk, the fatigue region of the blade was somewhat obscured by tan colored material consistent with "engine dust" and darker unidentified material. Blade 23 was removed from the disk and ultrasonically cleaned in acetone for further examinations.
The oxidized area of the blade showed three regions with differing fracture features. In the initial region, the fracture was relatively smooth and flat with marking typical of fatigue propagation in nickel based alloys. The area adjacent to the smooth region displayed a much rougher topography with some crack arrest marks. To the right (forward) of the rough region, the fracture was relatively clean with little oxidation, and contained features typical of overstress separations in cast high temperature alloys.
High magnification optical examinations revealed a shallow rounded pit on the trailing fir tree surface at the fatigue origin. The pit and fatigue initiation area were just outboard of the blade-to-disk contact area indicated by the fretting pattern on the blade.
The fracture face of blade 23 was further tape replica cleaned and examined on the scanning electron microscope (SEM). High magnification viewing uncovered finely spaced and oxidized fatigue striations throughout the initial fatigue region. Ratchet marks and striation orientations established that fatigue initiation was at multiple origins within the pit area. However, fine details indicated that the actual origins were consumed by the pit formation. Therefore, it could not be determined if the pit preceded the fatigue crack.
SEM viewing of the intermediate fracture region found mixed oxidized features. Many areas showed clear fatigue striations while adjacent areas contained heavily oxidized overstress features. The region forward of the second area of fracture surface showed clear unoxidized overstress features, typical of cast alloys.
Energy dispersive x-ray spectra acquired from the overstress region were consistent with the specified material, Inconel 100 (IN 100).
SEM views of the pit at the origin show it measured about 0.015 inches wide (fore-aft) and 0.010 inches tall (inboard-outboard).
During initial examinations of the turbine disk assembly a crack was noted in blade 25 at the same location as the fatigue region of blade 23. The crack visually measured 0.075 inches across the aft face and 0.10 inches along the side when removed from the disk. The crack intersected a corrosion/oxidation pit on the trailing side of the blade adjacent to the aft corner. The pit was very similar in characteristics to the pit on blade 23. Measurements found the pit about 0.030 inches wide and 0.040 inches tall.
The crack in blade 25 was forcibly opened in the laboratory after the blade was back cut. Opening revealed heavily oxidized and darkened crack surfaces. Fracture features were similar to the initial fatigue region on blade 23 with fatigue initiating at the pit at the aft trailing corner of the blade. Close examinations found that the fatigue in blade 25 initiated at multiple sites along the bottom of the pit subsequent to formation of the pit.
The remaining blades were removed from the disk and visually inspected. No additional cracks were found but a large number of blades exhibited single or multiple pits in similar locations to blades 23 and 25. Of the 27 total blades in the disk, all but 7 contained pits. Blade 26 was selected for metallographic sectioning.
Examinations of the polished specimen revealed general corrosion attack of the blade. Dark material consistent with oxidation/corrosion products partially filled the pit. The attack was along a wide front and did not appear to be following particular microstructural features. At its greatest, the pit measured about 0.004- to 0.005-inches deep. In addition to the corrosion attack, a transgranular crack (consistent with a fatigue crack) was uncovered in the pit. The crack measured about 0.006 inches long.
The complete Materials Laboratory report is contained in the public docket of this investigation.
The helicopter was equipped with an on-board video recording system, intended to provide customers with a video record of their tour flight. The Vehicle Recorder Division of the Safety Board received a Stack System 800 Digital Video Recorder module with 4GB compact flash memory card and a VHS tape recorded on board the accident flight.
At 43 minutes 50 seconds from the beginning of the video, while pointing out a waterfall to passengers, the pilot states that they are at, "a little over 3,000 feet." Approximately 44 minutes 44 seconds from the beginning of the recording, while filming from the left view camera, what appears to be a yaw motion to the right can be observed on the video. Immediately following the yaw, the recording changes to the internal camera view. An aircraft oscillation can be observed, and quickly increases in magnitude. Treetops become visible through the right window, and the passengers and pilot start leaning to the left. The video ends at approximately 44 minutes 55 seconds as the helicopter appears to contact the treetops. No engine noises can be heard on the video, and the pilot does not make any comments or utterances on the recording either just prior to, or during the descent.
The complete On Board Video Recording Factual Report is contained in the public docket of this investigation.
Results from the investigation
Turbomeca issued Service Letter No. 2436/06/ARL/182, dated April 12, 2006, notifying operators that Turbomeca was reducing the service life of the Arriel 1B second stage turbine blades from 6,000 hours to 3,000 hours.
Turbomeca issued Mandatory Service Bulletin No. A292 72 0807 that directed the inspection and established criteria for replacement of the second stage turbine wheel or blades of Arriel 1B engines.
The FAA issued Airworthiness Directive 2006-02-08R1 that defined Arriel 1B, 1D, and 1D1 second stage turbine blade inspection and replacement.
The FAA issued Airworthiness Directive 2006-06-17 that required an inspection of the nozzle guide vanes of the second stage turbine on Arriel 1B, 1D, and 1D1 engines modified to the TU 202 standard.
The Safety Board IIC released the aircraft wreckage to the owner/operator on January 26, 2006.