N206KKBell Helicopter Textron 206B · 1991

On October 11, 2013, about 1244 Pacific daylight time, a Bell 206L-1 helicopter, N206KK, was destroyed during a forced landing following an in-flight loss of engine power and fire during cruise flight near Seeley, California. The helicopter was registered to and operated by Blackhawk Helicopters Inc., El Cajon, California, under the provisions of Title 14 Code of Federal Regulations Part 91. The airline transport rated pilot and his two passengers were not injured. Visual meteorological conditions prevailed and a company visual flight rules flight plan was filed for the aerial observation flight. The cross-country flight originated from Yuma, Arizona, about 1200 with an intended destination of El Cajon. In a written statement to the National Transportation Safety Board (NTSB) investigator-in-charge (IIC), the pilot reported that during cruise flight, he heard a loud pop followed by an immediate loss of engine power. The pilot lowered the collective, and entered an auto rotation. He further reported that the left front door opened, and the cabin became filled with black smoke. Subsequently, the helicopter landed upright in an open area, and the pilot and passengers egressed. The pilot stated that upon exit, he saw fire originating from about 2 feet aft of the right forward door.

Examination of the helicopter by a Federal Aviation Administration (FAA) inspector revealed that the fuselage was mostly consumed by fire. All primary components of the helicopter were located at the accident site. The wreckage was recovered to a secure location for further examination.

The wreckage was examined at the operator's hangar October 21, 2013 by representatives from Bell Helicopter, Rolls-Royce, Blackhawk Helicopters, and the Federal Aviation Administration.

Examination of the wreckage revealed that little of the aluminum structure, magnesium components, or plastic/composite material remained. All cockpit instruments and controls were consumed by fire. An approximate 4 foot section of the tail boom, including the tail rotor and vertical fin, was not damaged by fire. The tail rotor turned and changed pitch freely, and exhibited no evidence of impact damage or operational failure.

The main rotor blades had been largely consumed by fire. The blade grips were present on the main rotor mast, and were fire damaged. The main rotor transmission housing was consumed by fire, exposing the transmission gearing within. A few feet of each main rotor blade's tip were also present, and exhibited fire damage. The leading edges exhibited no evidence of impact damage.

The engine exhibited evidence of exposure to an intense fire, and the magnesium auxiliary gearbox housing was consumed by fire. The engine also exhibited evidence of high-energy debris being liberated from the turbine section. Several gears from the accessory gearbox were found within the debris, however not all the gears could be accounted for.

The helical torquemeter gear and power output gear were found in the debris; however, the idler gear and N2 tachometer gear were not located. The centrifugal breather gear and two unidentifiable idler gears were found melted within a large pool of solidified aluminum. The fuel pump/oil pump gear was found, and was thermally damaged. The spur adapter gearshaft remained in-place, and was thermally damaged. No engine accessories were located in the recovered debris.

Evidence of high-energy debris exiting the turbine was observed in the engine exhaust collector. Visual examination of the turbine revealed the 4th stage turbine wheel turbine blades were separated at the blade roots. A boroscope was inserted into the power turbine module, which revealed several missing 3rd stage turbine blades.

The leading edges of the compressor blades were rough to the touch, and exhibited evidence of exposure to intense heat. Otherwise, inspection of the compressor inlet revealed no visible damage or signs of foreign object ingestion.

The engine and two previously removed engine oil filters were subsequently shipped to Rolls-Royce for further examination.

Engine Examination

On November 13, 2013, the engine was examined under the supervision of the NTSB IIC at Rolls-Royce's facility near Indianapolis, Indiana.

Disassembly of the engine revealed that the exhaust collector and exhaust duct exhibited impact marks and holes consistent with high-energy debris. There was also evidence that high-energy debris penetrated the starboard-side compressor discharge tube. Several of the exit holes on the exhaust collector were in-line with gearbox-mounted accessories, including the fuel control unit and power turbine governor. The compressor module was then disassembled.

The 2 1/2 bearing was visually inspected, and no evidence of operational damage was noted. The compressor front support and number 1 bearing housing was removed. The number 1 bearing was dry, and exhibited no evidence of operational damage. The compressor's number 2 bearing was removed, and was found dry; however, it exhibited no evidence of operational damage. The compressor bleed valve housing and diaphragm had been consumed by fire. The remains of the valve spring and poppet valve exhibited no evidence of operational damage. The compressor shroud housing was removed, which exposed the impeller. No evidence of foreign object damage was noted, and no evidence of the impeller contacting the shroud during operation was present.

The helical power train drive assembly, which was the only accessory gearbox component that was found intact, was removed and examined. The assembly consisted of the housing, pinion gear, and number 3 and number 4 bearings. The pinion gear coupling was bound within the pinion gear. The number 3 bearing was intact, dry, and coated with powdery residue. The separator cage of the number 4 bearing was fractured, and rotational scoring on the outer surface of the bearing race was observed. The number 4 bearing rollers appeared to have been exposed to intense heat, with one of the rollers exhibiting evidence of operational damage. The damaged roller was flattened and out-of-round.

The number 5 bearing retaining snap ring was found dislodged. The outer race exhibited impact damage. Following removal of the number 5 bearing, two fragments of the race liberated, which exposed the ball bearings. The ball bearings did not exhibit any evidence of operational damage.

The turbine to compressor coupling was removed, and found fractured in two locations, with the forward end of the coupling stuck within the power turbine outer shaft, and the aft end jammed within the power turbine rotating labyrinth seal. The fracture surfaces were found consistent with overload.

The combustion outer case and inner liner were removed. Neither exhibited evidence of failure or abnormal combustion. Removal of the combustion section exposed the 1st stage turbine nozzle shield. The shield appeared to be undamaged, and did not exhibit any evidence of any metal splatter. The number 8 bearing sump cover exhibited evidence of contact with the turbine tie bolt and lock nut. The number 8 bearing was dry; however, it exhibited no evidence of any operational failure. The bearing lock nut was found backed off and loose within the bearing chamber. Impact marks on the underside of the number 8 bearing sump cover are consistent with the lock nut being centered and in place on the tie bolt. The tie bolt was found fractured. Both fragments were removed, and exhibited signatures consistent with tension overload.

The oil supply fitting for the 6/7 bearing and the number 8 bearing were coated with blackened oil, all passages and filter screens were clear. The number 6 and number 7 bearings were dry, and exhibited no evidence of operational failure.

The exhaust collector was then separated from the turbine module. The nozzle exhibited impact damage to the trailing edges of several vanes. The nozzle's inner web had separated and had been deformed from contact with the power turbine outer shaft. The 4th stage turbine rotor blades and shroud were separated. The fracture signatures were consistent with overload. The rotor's aft curvic coupling was smeared, and the forward face of the rotor's web exhibited rotational scoring.

The 3rd stage turbine rotor shroud was separated, and all of the turbine blades were separated about mid span. The separation surfaces were found consistent with overload. The 3rd stage turbine nozzle exhibited impact damage on several vanes. The forward curvic coupling was smeared.

The 2nd stage turbine rotor was intact, but the stub shaft portion of the turbine wheel had fractured. The stub shaft was determined to have failed in overload. The turbine blades exhibited impact marks on the trailing edges of several blades.

The 1st stage turbine rotor was intact and exhibited impact and rub damage to the leading edges of all the blades. The 1st stage turbine nozzle exhibited impact damage along the trailing edges of several vanes.

Metallurgical Examination

Examination of various components was conducted by Rolls-Royce's Failure Analysis Laboratory with permission from the NTSB Material Laboratory.

The number 4 bearing, power train pinion gear, and turbine shaft to pinion gear coupling were oxidized from the engine fire. The turbine shaft to pinion gear coupling rotated in the power train pinion gear consistent with spline disengagement. The number 4 bearing separator was fractured through one roller pocket at three locations. The separator fragment was retained within the bearing. The surface of the rail fractures were oxidized, and a specific fracture mode could not be identified. All but one of the rollers were cylindrical in shape, and showed no evidence of skidding damage. However, a roller from the fractured pocket exhibited a flat spot, which was consistent with skidding damage.

The outer ring raceway was distorted, and the rail surfaces were worn from contact with the separator after the rails fractured. The raceway was distorted, and the inner diameter surface was severely worn. The wear on the inner diameter surface was consistent with the inner ring having spun on the power train pinion gear. Both rings were oxidized from the post landing engine fire.

A cross-section through the number 4 bearing inner ring revealed that it was worn, distorted, and radial cracks emanated from the raceway surface, and extended radially inward. The entire microstructure was thermally damaged, consistent with damage generated from spinning on the pinion gear and the post landing fire. Cross-sections through the number 4 bearing outer ring and the flat spotted roller were worn and distorted, which is consistent with a skidding roller. The entire microstructures were also thermally damaged. Semi-quantitative x-ray fluorescence (XRF) analysis determined that the outer ring and a roller were an AMS 6491 (M-50 steel) type material, and the inner ring was an AMS 6278 (M-50 NIL steel) type material as required by the engineering drawing. In addition, the number 4 bearings showed evidence of spinning on the Power Train Pinion Gear.

According to the Rolls-Royce representative, the generated heat caused the expansion and softening of the aft end of the Power Train Pinion Gear, resulting in the power train pinion gear disconnecting from the turbine shaft-to-pinion gear coupling splines. As a result of this disconnect, the load on the power turbine was suddenly removed, which resulted in an over speed of the 3rd and 4th stage turbine wheels and subsequent separation of their shrouds and blades, some of which were liberated from the engine.

In addition to the turbine components, the oil filters, which had been retained by the operator's mechanic during the previous troubleshooting, were examined. The oil filters were back-flushed and the particle contaminants analyzed. The chemical makeup of the particles was consistent with two different alloys, AMS9310 low alloy steel and M50 NIL. The Pinion Gear is manufactured from AMS9310, and M50 NIL is material used within the bearing races.

Maintenance Records

Review of the helicopter maintenance records revealed that in June, 1996, the gearbox logbooks depicted compliance with CEB 72-3193 Rev-2 by Omni Helicopters (while the gearbox was installed on a different engine, CAE-895018). Incorporation of this CEB installed both the number 3 and number 4 bearings, part number 23030917. The gearbox logbooks specified that the number 4 bearing, serial number JJ-01236, was installed at this time, along with #3 bearing, serial number JJ-01235.

In October, 2000, the gearbox was subjected to a sudden stoppage/hard landing inspection while owned by Omni Helicopters. The gearbox transferred ownership from Omni Helicopter to JBI Aviation in 2001. JBI Aviation sent the gearbox to Arrow Aviation for the required sudden stoppage/hard-landing inspection. During this inspection at Arrow Aviation, the logbooks indicated that CEB 72-3193 Rev-3 was complied with. However, at the time of this inspection, CEB A-72-3217 should have been incorporated, which requires the replacement of the number 3 and 4 bearings with the new part number, 23066678. There was no indication within the logbooks that any bearings were changed at this time.

In April 2010, the turbine module was overhauled at AeroMaritime. The work order paperwork specified that no CEBs were accomplished during the overhaul. However, an entry within the gearbox logbook, dated April 17, 2010, specified the number 5 bearing in the gearbox was replaced in order to comply with CEB 72-3158. There was no mention in the logbooks of CEB A-72-3217 being complied with at any time.

During engine disassembly, the number 3 bearing was found to have been updated with the new part, to part number 23066678. It could not be determined from the logbooks when this bearing had been replaced with the updated part. However, the number 4 bearing found installed on the accident engine had the same serial number and part number that was installed in 1996.

CEB-A-72-3217 specifies that all Rolls-Royce Model 250-C30/C30S engine, serial numbers CAE-890001 though 890888, must have the #3 and #4 Bearings replaced with new part number 23066678 no later than 31 December 2002. This CEB was released on 15 January 1998, with the latest Revision (4) released on 4 October 2001.