Abilene, TX, USA
N708M
BELL 206L-3
Following the post-flight inspection of a repositioning flight, the pilot noted a crack in the trailing edge of the rotor blade. He had characterized the previous flight as turbulent, but otherwise unremarkable and without any mechanical anomalies. An examination of the main rotor blade revealed a path of interconnected porosity between the lower trailing edge skin, trailing edge strip, and epoxy that allowed for the introduction of moisture and resultant corrosion. The failure was due to insufficient squeeze-out of the epoxy layer between the lower skin and the trailing edge strip during manufacture. The corrosion pit allowed the formation of a fatigue crack with a resultant crack along the blade surface.
On March 29, 2009, approximately 1820 central daylight time, a crack was discovered on one main rotor blade on a Bell 206 L-3, N708M, operated by Air Evac EMS, Inc. The helicopter had just returned from a mission and landed uneventfully at the operations base in Abilene, Texas. Visual meteorological conditions prevailed during the mission and repositioning flight. The repositioning portion of the flight was being conducted under the provisions of Title 14 Code of Federal Regulations Part 91 without a flight plan. The pilot and two medical crewmembers were not injured. According to the operator, the preflight for the mission was without anomalies or issues. The mission flight and repositioning flight were turbulent, but otherwise, unremarkable. During the post-flight inspection of the helicopter, the pilot discovered a crack in one of the main rotor blades. The pilot states specifically that he was unable to detect any vibration in the system due to the turbulence. The operator removed the blade and sent it to the National Transportation Safety Board Materials Laboratory in Washington, DC. On July 15, 2009, an investigator with the Safety Board and an engineer from Bell Helicopter examined the blade. The crack was observed in the top and bottom skin of the trailing edge of the blade. The crack ran chord-wise approximately 8.75 inches from the trailing edge of the blade towards the leading edge of the blade. The crack then branched into two cracks running longitudinally in opposite directions; one crack measuring 9.5 inches inboard and the second crack measuring 7.75 inches outboard. Examination of the crack surface revealed features consistent with a fatigue crack that initiated from a pit along the leading edge of the trailing edge strip. The crack’s initiation site revealed a yellow discoloration consistent with corrosion. Further examination revealed a region of interconnected porosity that formed a channel through the adhesive layer that bonded the lower skin to the trailing edge strip. According to Bell Helicopter, during the manufacturing process, the blade is taken through two separate leak checks designed to detect porosity in the epoxy that bonds the upper and lower skins to the trailing edge strip. According to assembly records, these two leak checks were conducted by October 23, 2005, and October 26, 2005, respectively. There were no notes in the assembly records to indicate that a leak was noted; however, records did indicate, through a quality assurance stamp, that the leak checks had been performed. According to Bell Helicopter, if a leak is noted, additional adhesive is applied and an entry in the assembly records is not necessarily made. Engineers and investigators from Bell Helicopter reported that trailing edge blade cracks are a rare occurrence. If a trailing edge blade crack is not found during normal maintenance and pilot checks, then the crack will eventually become large enough that the pilot is able to detect a one per revolution vibration in flight. They stated that when these events occur, the pilot is able to easily detect an imbalance situation in the operation of the aircraft; the situation does not propagate beyond the imbalance and vibration. The cracks are usually noted during the post flight inspection, while trying to diagnose the imbalance. According to the operator and company maintenance records, they had experiences a one per revolution vibration and had conducted tracking and balancing with the blade, in order to track the vibration. It was noted that after this maintenance, the vibration was no longer noted.
The formation and propagation of a fatigue crack in the trailing edge of the main rotor blade due to interconnected porosity and resultant corrosion. The area of interconnected porosity was due to a manufacturing defect which was not detected during the manufacturing process.
Source: NTSB Aviation Accident Database
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