SEA06TA153

The pilot reported that the helicopter's engine failed during performance of a power assurance check. Prior to beginning the power check, the pilot established straight and level flight at 90 knots and 8,000 feet msl (about 5,000 feet agl). He increased collective, the helicopter began to climb, and at approximately 8,300 to 8,400 feet msl and a torque reading of 95%, there was a loud grinding noise and the engine lost power. The pilot entered an autorotation. During touchdown in an open field, the main rotor struck and severed the tail boom. Examination of the engine determined that the turbine exhibited internal damage and would not rotate. Disassembly of the turbine section revealed rub damage to all the blade tips on the first and second stage turbine wheels. Corresponding rub damage was noted to the blade paths in the second stage turbine nozzle for these turbine wheels. A single blade was separated from the second stage turbine wheel. There was severe damage to the turbine wheels and nozzles located downstream of the second stage turbine wheel. Metallurgical examination found that the second stage turbine nozzle assembly second stage blade path had been dimensionally restored using thermal spray material. The coating exhibited unmelted particles, oxides, and was separated from the base metal at both forward and aft ends. The defects in the coating resulted in blade tip rub and led to the fatigue failure of a blade, which separated and destroyed the downstream turbine wheels and nozzles. Review of the helicopter's maintenance records indicated that the engine's turbine section was replaced with a repaired unit approximately 13 flight hours prior to the accident. During the repair of the turbine, dimensional restoration of the second stage turbine nozzle second stage blade path was performed.

On August 3, 2006, about 1230 mountain daylight time, a Bell 206L-1, N5746N, landed hard during an emergency landing following a loss of engine power near Busby, Montana. The commercial pilot and the three passengers were not injured. The helicopter, which was registered to TS Aviation of Worland, Wyoming, dba Sky Aviation, and under the operational control of the Bureau of Indian Affairs (BIA), U.S. Department of the Interior (DOI), sustained substantial damage. Visual meteorological conditions prevailed and a company flight plan was filed for the local public use flight. The purpose of the flight was to perform aerial wildfire reconnaissance.

The pilot reported to a DOI investigator that the engine failed during performance of a power assurance check. Prior to beginning the power check, the pilot established straight and level flight at 90 knots and 8,000 feet msl (about 5,000 feet agl). He increased collective, the helicopter began to climb, and at approximately 8,300 to 8,400 feet msl and a torque reading of 95%, there was a loud grinding noise and the engine lost power. The pilot entered an autorotation. He had difficulty stabilizing the rotor RPM during the initial stages of the autorotation, but had established a steady-state autorotation prior to descending below 500 feet agl. During touchdown in an open field, the main rotor struck and severed the tail boom.

Examination of photos taken at the accident site by the DOI investigator revealed that both of the cross tubes were folded aft, and the fuselage of the helicopter was resting on the ground.

The helicopter was powered by a Rolls-Royce Model 250-C30P gas turbine engine, S/N CAE 890629. Initial inspection of the engine by the DOI investigator and operator personnel revealed that the turbine exhibited internal damage and would not rotate. Outward dents and exit wounds were noted on the turbine exhaust collector and the engine cowling. The engine was removed by the operator and sent to the facilities of Rolls-Royce Corporation in Indianapolis, Indiana for further examination.

The engine was examined at the facilities of Rolls-Royce Corporation in Indianapolis, Indiana, on September 7 and 8, 2006, under the supervision of the DOI investigator. Disassembly of the turbine section revealed rub damage to all the blade tips on the first and second stage turbine wheels. Corresponding rub damage was noted to the blade paths in the second stage turbine nozzle for these turbine wheels. A single blade was separated from the second stage turbine wheel. There was severe damage to the turbine wheels and nozzles located downstream of the second stage turbine wheel. Metallurgical examination of the second stage turbine wheel and nozzle by Rolls-Royce personnel determined the following:

1. The separated second stage turbine wheel blade cracked in fatigue. No anomalies were found at the origin site.

2. Two additional second stage turbine wheel blades cracked in fatigue but did not separate.

3. The microstructure, hardness, and chemistry of the second stage turbine wheel conformed to the requirements of the engineering drawing.

4. The second stage turbine nozzle assembly second stage blade path had been dimensionally restored using thermal spray material. The coating exhibited unmelted particles, oxides, and was separated from the base metal at both forward and aft ends.

Review of the helicopter's maintenance records indicated that the engine received a 100-hour inspection and turbine replacement on July 27, 2006, approximately 13 flight hours prior to the accident. The turbine installed on July 27, 2006 was identified as P/N 23031493, S/N CAT 98201. The FAA Form 8130-3, Airworthiness Approval Tag, for the turbine identified it as being repaired by Rolls-Royce Engine Services - Oakland Inc. and was dated March 2, 2006. According to Rolls-Royce personnel, dimensional restoration of the second stage turbine nozzle second stage blade path was performed at Rolls-Royce Engine Services - Oakland Inc. as part of the repair of the turbine.