On August 28, 2013, about 1030 mountain standard time, a Bell Helicopter Textron 206L-1, N2018R, sustained substantial damage during an off airport emergency landing, about 14 miles west of the Payson Airport (PAN) Payson, Arizona. The helicopter was registered to GM leasing Company LLC and operated by Airlift Helicopters under the provisions of Title 14 Code of Federal Regulations Part 135. The commercial pilot sustained serious injuries and none of the 5 passengers were injured. Visual meteorological conditions prevailed and a company visual flight rules, flight plan was filed for the cross county corporate flight that departed Scottsdale Airport, Scottsdale, Arizona at 0910 with a planned destination of Flagstaff Pulliam Airport, Flagstaff, Arizona.
The pilot reported a partial loss of engine power while in cruise flight at about 1,500 feet, above ground level, and performed an autorotation emergency landing. He maneuvered the helicopter during descent to avoid obstacles and said he may have gotten slow. The pilot also mentioned that a low rotor warning light was illuminated during the descent along with the N1 indication dropping to about 70 percent. Subsequently, the helicopter landed hard and the tail boom separated.
The helicopter was recovered to a local storage facility for further examination.
Initial examination of the helicopter airframe revealed that main rotor flight control continuity was established from the cockpit controls to the main rotor pitch links. Tail rotor flight continuity was established from the left anti-torque pedal assembly to the separation of the tail boom at the aft fuselage. Tail rotor flight control continuity could not be established from the pilot's right-hand side anti-torque pedals due to a fractured control tube from the pedals to the center console. However, continuity from the fractured ends of the control tube was established.
Drive continuity was confirmed from the main rotor system down through the transmission and the engine to the tail rotor driveshaft, at the separation of the tail boom at the aft fuselage. Both main rotor pitch change links were observed fractured. Tail rotor drive continuity was confirmed from the tail rotor forward to the separation of the tailboom at the aft fuselage. The tail rotor was rotated by hand and rotated freely in both directions.
The landing skid appeared spread and the cross-tubes were rotated aft. The fuselage, doors, windows and the main cabin were intact.
When electrical power was supplied to the instrument system, through the helicopter battery, the engine over-temp light illuminated.
No anomalies were observed with the flight, navigation, or propulsion instruments.
The engine remained attached to the airframe and no visible damage was observed. The power turbine governor was examined and correct responses were observed with movement of the collective. Rotation of the throttle resulted in proper movement of the fuel control throttle arm. Manual rotation of both the N1 and N2 drive train was accomplished. The engine was removed from the airframe for shipment to the manufacturer. The helicopter's original Rolls Royce 250-C28B engine had been replaced with a higher horsepower Rolls Royce 250-C30P engine.
Examination of the engine at the facilities of Rolls Royce indicated that the engine could be safely test run. During the engine run, following the start, erratic and higher than manufacturer specified turbine vibrations were present. During engine operation it was determined that the engine was not capable of making specified power. Further, the engine responses indicated a possible power turbine governor malfunction and therefore the governor was replaced with a new original equipment manufacturer (OEM) governor, but the test results were the same. It was determined that further engine run attempts would not be accomplished and that the engine would be disassembled and further examined.
During engine disassembly, the outer portions of the first stage turbine blades were observed to have sustained thermal damage at all blade tips in varying degrees. The stage one nozzle sustained thermal damage to the trailing edges of three vanes, where localized melting was observed. A metallurgical examination of the turbine blade's microstructure revealed localized metal temperatures in excess of 1149 degrees Celsius. According to the airframe manufacturer's Rotorcraft Flight Manual, the maximum continuous limitation for the Turbine Outlet Temperature (TOT) is 716 degrees Celsius and a maximum takeoff TOT of 768 degrees. The investigation was unable to determine when the over-temperature occurred. Blade examination revealed that the hardness and composition of the first stage turbine wheel met engineering specifications. Further, there was no evidence of fatigue progression observed on any of the fractures.
Additional testing was accomplished on the associated components. The power turbine governor, and fuel control unit were tested at the manufacturer and both units met acceptance testing specifications. The fuel nozzle flow was flow tested and met specification. The thermocouple harness met testing specifications. Further, the MGT gauge was Barfield tested and met specification criteria.
The postaccident examination of the airframe and engine revealed no additional evidence of a mechanical malfunction that would have precluded normal operation.
The maximum gross weight limitation for flight in the helicopter, per Bell's Rotorcraft Flight Manual is 4,150 pounds. The pilot stated a gross weight of 4,000 pounds on the National Transportation Safety Board (NTSB) pilot/operator accident/incident report at the time of the accident. A review of the fuel burn charts, calculated the fuel burnt since takeoff was about 210 pounds. Adding this fuel to the helicopter's reported weight by the pilot at the time of the accident, would compute the takeoff gross weight to be about 4,210 pounds. Further, the take-off gross weight was calculated by the NTSB investigator-in-charge (IIC), using the empty weight of the helicopter, the reported weight of the occupants, 70 gallons of fuel, and 21 pounds of baggage. The calculated gross weight at the time of departure was about 4,222 pounds, which exceed the manufacturer's maximum gross weight limitation.
Utilizing the weather conditions at the nearest reporting station, the density altitude was calculated by the IIC to be about 9,500 feet for the cruise altitude of the accident flight.