N405PH
Bell 407
The helicopter, which had a FADEC controlled turboshaft engine installed, experienced a loss of engine power after take off from an offshore platform in the Gulf of Mexico. The 10,000-hour commercial rated helicopter pilot, reported that the helicopter "cleared" the platform deck and entered translational lift when "it lurched violently to the left and started a rapid descent." The pilot then lowered the collective and dropped the nose to avoid striking the platform. He recalled hearing a "screaming or screeching type of sound" as the helicopter descended, and was not able to maintain main rotor RPM. The helicopter landed into the water, and rolled to the right. After the occupants evacuated, the helicopter submerged 30 feet under the surface. The skid mounted float system was not activated, although the pilot stated that the floats were "armed" and that he attempted to deploy the floats while in the autorotation, via the float activation button that was mounted on the collective. The float activation button was located within a circular ring that was mounted on the pilot's collective. Download of the Incident Recorder (IR) of the ECU showed that the IR was first triggered by a droop in rotor RPM, followed by a MGT exceedence, power surge, engine flameout, and Ng underspeed in a span of 3.38 seconds. Download of the MGT instrument showed a recorded peak temperature of 953 degrees for 1 second. Teardown examination of the engine revealed a catastrophic failure of the power turbine assembly. Metallurgical examinations by the NTSB and Rolls Royce revealed evidence that the 3rd stage turbine wheel airfoil(s) had failed. A root cause of the failure was not determined, and further testing was conducted at Rolls Royce. As a result of some of the dynamic test results, Rolls Royce issued several Commercial Engine Bulletins (CEBs) regarding inspection of in-service turbine assemblies. Additionally, Bell Helicopter issued an Alert Service Bulletin (ASB), which introduced flight manual revisions to avoid power turbine RPM (Np) steady state operation between 68% and 97%. After examinations of the airframe (fuel system, rotor systems, flight control systems, drive systems, electronic control systems), no anomalies were found other than within the power turbine assembly.
On November 16, 2003, approximately 1035 central standard time, a Bell 407 helicopter, N405PH, registered to and operated Petroleum Helicopters Inc. (PHI), of Lafayette, Louisiana, submerged following an autorotation into open ocean water following a loss of power after takeoff from an offshore platform (Eugene Island 27) in the Gulf of Mexico. The commercial pilot and his two passengers evacuated the helicopter after it entered the water, and were not injured. Visual meteorological conditions prevailed and a company VFR flight plan was filed for the Title 14 Code of Federal Regulations Part 135 on-demand air taxi flight. The flight was originating at the time of the accident and its intended destination was an offshore platform, Eugene Island 63. The 10,000-hour commercial rated helicopter pilot, 1,200 hours of which were in the Bell 407, reported that the helicopter "cleared" the platform deck and entered translational lift when "it lurched violently to the left and started a rapid decent." The pilot then lowered the collective and dropped the nose to avoid striking the platform. He recalled hearing a "screaming or screeching type of sound" as the helicopter descended, and was not able to maintain main rotor RPM. The aircraft landed into the water, and rolled to the right. After the occupants evacuated, the helicopter submerged 30 feet under the surface. The skid mounted float system was not activated, although the pilot stated that the floats were "armed" and that he attempted to deploy the floats while in the autorotation, via the float activation button that was mounted on the collective. The float activation button was located within a circular ring that was mounted on the pilot's collective. The helicopter, serial number 53207, had accumulated a total time of 4,606.7 hours. Research of the airframe records showed the last major inspection accomplished was the 300-hour/3-month inspection at 4,482.3 total airframe time on August 26, 2003. All of the life limited components were within limits and verified. Part numbers and serial numbers of actual installed components was verified. Airworthiness Directive (AD) research showed that all directives were accomplished prior to their required due date. The engine, a Rolls Royce model 250-C47B, serial number CAE-847473, had accumulated a total time of 2,591.9 hours. Engine records showed the last major engine inspection was the 150-hour inspection accomplished at 2,456.4 total engine hours, on August 19, 2003. In addition, the 300-hour inspection was accomplished at 2325.4 total engine hours, on June 1, 2003. According to records, the turbine assembly p/n-23063354, s/n-CAT44132, was last overhauled on November 7, 2002. Total operating time of the assembly was 5,870.1 hours, and 972.9 hours remaining until the next overhaul. Critical turbine component time (comprised of 4 turbine wheels) as of November 16, 2003, were: 1st stage wheel - 706.35 hours; 2nd stage wheel - 706.35 hours; 3rd stage wheel - 2,586.05 hours; 4th stage wheel - 2,586.05 hours. Recovery: After the helicopter was raised from the water and placed on a recovery vessel, it was washed with fresh water. It was noted that the skid mounted emergency floats were not deployed, and both float inflation bottles were fully charged. The float inflation bottles were discharged, and the skid landing gear and main rotor blades were documented and removed to facilitate transport. The Electronic Control Unit (ECU) was removed, preserved and shipped to Goodrich, West Hartford, Connecticut, for data retrieval. The Ng, torque, and MGT solid state indicators were removed and shipped to MOOG Components Group, Springfield, Pennsylvania, for download. The following observations were noted during examination at the PHI facility located in Lafayette, Louisiana: Airframe: Except for corrosion from salt-water emersion, the airframe exhibited minor damage. The interior of the fuselage/cabin was intact and all seats and occupant restraint systems were found attached and secure. Flight control continuity was confirmed from the cockpit to the hydraulic servos (collective, and left/right lateral) up to the swashplate mixing unit. The float-arm switch was found in the ARMED position and the throttle twist grip was found positioned slightly below the MAX position. Four of the six skid mounted float bags were contained in their respective enclosures, and the remaining two float bags were out of their enclosures, but had not been inflated. Electrical continuity checks were conducted on the float inflation circuits from the float activation button to the inflation bottles, and no anomalies were discovered. The tail boom was still attached to the fuselage and was not damaged, and the lower portion of the right horizontal stabilizer end-plate was collapsed to the left from overload. Hydro-dynamic deformation was observed on the leading edge of the lower portion of the vertical fin. Drive Systems: The main driveshaft was intact and attached to the engine and transmission. The main transmission could not be rotated, however, recovery personnel reported that the transmission rotated freely when the helicopter was raised from the water. Tail rotor drive continuity was confirmed throughout the shaft system. Rotor Systems: The main rotor blades had been removed from the hub during recovery. Although two of the main rotor blades exhibited chordwise overload fractures, no rotational damage was noted on any of the four blades. The tail rotor assembly was intact and attached to the tailboom and was not damaged. Engine: After recovery, visual inspection showed some salt-water corrosion. N1 and N2 were found locked and would not rotate. All rigid and flexible oil and fuel lines were intact and all fittings were found to be tight. When removed from the outer combustion case, the fuel nozzle inlet line contained fuel at the nozzle. The nozzle was corroded, and the condition of the fuel nozzle screen could not be assessed. The engine oil filter canister was not damaged and the filter element was free of visible debris. The right side of the gas producer turbine support was fractured from the 3-o'clock position on the aft flange toward the 12-o'clock position on the forward flange. The inner surface of the right compressor air discharge tube exhibited dents adjacent to the fracture in the gas producer turbine support. The exhaust collector support and exhaust duct assembly exhibited numerous dents on their surfaces, and the lower right forward side of the exhaust collector support had a tear at the 5-o'clock position. The NTSB IIC determined, and parties concurred, that the engine assembly required a complete teardown examination. Engine Teardown Examination: When the turbine exhaust collector support was removed from the accessory gearbox, the #5 bearing retaining ring and spacer were found dislodged. Pieces of what appeared to be turbine wheel hub and blade pieces were found on top of the 4th stage nozzle, and the 4th stage nozzle was scarred. there was no evidence of heat distress of the #5 bearing, and the bearing nut was secure. The spur adapter gearshaft was not damaged. The 1st stage turbine wheel was found fractured into five pieces. The turbine to compressor coupling was fractured approximately 3-inches forward of the turbine end, and a section of the tie bolt remained lodged in the 3-inch section of the coupling. The longer segment of the bolt was not located. Second stage nozzle vanes were missing in a 200-degree arc from the 11-o'clock position clockwise to the 7-o'clock position. The 3rd stage nozzle turbine nozzle exhibited heavy scarring and was fractured at the 3-o'clock position and 9-o'clock position. When the 2nd stage turbine wheel was removed, the diaphragm of the 3rd stage nozzle and the hub of the power turbine support were lodged on the 2nd stage turbine wheel shaft. The struts that supported the hub in the center of the power turbine support were found sheared nearly flush with the hub. The outer nut that secured the outer turbine shaft to the 3rd stage turbine wheel was found properly torqued and secured. All of the airfoils from the 3rd stage turbine wheel were missing and appeared sheared from the wheel hub near their respective bases. ECU Download Data / Ng, Torque, and MGT Instrument Data: Download of the Incident Recorder (IR) of the ECU showed that the IR was first triggered by a droop in rotor RPM, followed by a MGT exceedence, power surge, engine flameout, and Ng underspeed in a span of 3.38 seconds. Download of the MGT instrument showed a recorded peak temperature of 953 degrees for 1 second. Metallurgical examinations by the NTSB and Rolls Royce revealed evidence that the 3rd stage turbine wheel airfoil(s) had failed. A root cause of the failure was not determined, and further testing was conducted at Rolls Royce. As a result of some of the dynamic test results, Rolls Royce issued several Commercial Engine Bulletins (CEBs) regarding inspection of in-service turbine assemblies. Additionally, Bell Helicopter issued an Alert Service Bulletin (ASB), which introduced flight manual revisions to avoid power turbine RPM (Np) steady state operation between 68% and 97%.
The loss of engine power due to the failure of the 3rd stage turbine wheel and subsequent catastrophic failure of the turbine assembly.
Source: NTSB Aviation Accident Database
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