WPR18LA152

The pilot was performing external load operations in the helicopter when he felt a vibration during a left descending turn. The pilot increased the collective to arrest the rate of descent, but the vibration became more prominent. He subsequently performed a forced landing, during which the helicopter impacted the ground and a main rotor blade contacted and severed the tailboom. The pilot did not report any malfunctions with the engine nor any abnormal increase or decrease in main rotor rpm. Postaccident examination revealed a fractured drive link and a fractured strap pack assembly. Examination of the fractured surfaces revealed signatures consistent with tensile overstress, most likely from the blade contacting the left inlet fairing after the helicopter impacted the ground. There was no other evidence of preimpact mechanical malfunctions or failures that would have precluded normal operation, and the reason for the vibration could not be determined based on the available information.

On May 29, 2018, at 1045 mountain standard time, an MD Helicopters 369FF, N975BW, was substantially damaged when it was involved in an accident near Wikieup, Arizona. The pilot and the passenger were not injured. The helicopter was operated as a Title 14 Code of Federal Regulations Part 133 external load flight. The purpose of the flight was to assist power line construction. As the pilot was conducting a slow, left descending turn about 100 ft above ground level, the helicopter started to vibrate. The pilot increased the collective to arrest the rate of descent, but the vibration became more prominent. He then performed a forced landing; the helicopter impacted the ground, and a main rotor blade contacted the tailboom, which separated from the empennage. The pilot did not report any malfunctions with the engine nor any abnormal increase or decrease in main rotor rpm. The helicopter was manufactured in 1994 and had accumulated 6,713.5 hours total time at the time of the accident. Review of maintenance records indicated that the last 100-hour inspection was completed 63.1 hours before the accident flight. The last 300-hour inspection, main rotor damper attachment and weight loading extension check were completed 252.9 hours before the accident flight. The helicopter was equipped with a Rolls-Royce (Allison) 250-C30 turboshaft engine, serial number CAE-900135. The engine had accumulated 6,766.1 hours at the time of the accident. The last 150-hour engine inspection was performed 48.0 hours before the accident. Recovery personnel reported that all main rotor blades were attached to their respective pitch housings/lead-lag links and dampers by main rotor folding pins and attaching hardware. The airframe exhibited a small area of crushed and torn skin on the right side, and the lower fuselage skin was deformed below the left cabin door. The left and right landing gear struts and skid tubes appeared undamaged. The horizontal stabilizer, vertical stabilizers, and stinger exhibited no visible damage. All five main rotor blades exhibited trailing edge bond separation that was more pronounced in the first 3 ft outboard of the root fitting doublers and lessened along the blade span. One rotor blade exhibited blue/green paint transfer marks on the upper and lower surfaces on the leading-edge area near the root fitting and doubler. These paint marks corresponded to the blue/green paint found on the left inlet fairing. At the root end of the same blade, the leading edge and lower surface revealed paint transfer from contacting the upper wire strike arm mounted on the canopy frame. In addition, all laminates (straps) of the same blade strap assembly were fractured. The majority of the straps fractured about 5 inches inboard of the lead-lag bolt. The rotating swashplate scissor link was fractured, and the rotating swashplate was displaced in the lag direction. The upper hub displayed contact marks on the lag side from contact with the five pitch housings. Theelastomeric dampers on three blades appeared normal. The fourth blade's elastomeric damper was pulled out of the damper housing. The fifth blade's elastomeric damper was detached from the housing and extended out of the housing about 3.5 inches. Drive continuity from the engine driveshaft to the main and tail rotor driveshaft was verified when the shaft was manually rotated. Tail rotor driveshaft continuity was established from the main transmission outboard to the separation point at the tailboom. Both tail rotor blades were undamaged. The tail rotor control rod was continuous from where it separated in the tail boom to the bellcrank at the tail rotor transmission. Control continuity was established from the cyclic and collective controls to the stationary swashplate. Anti-torque control continuity was verified from the anti-torque pedals to the separation point at the tailboom. Additionally, continuity was confirmed from the separated tailboom section to the tail rotor assembly pitch control. The fractured drive link, fractured strap pack assembly, and parts of engine air inlet fairing were examined at the NTSB Materials Laboratory. The examination revealed that the drive link attachment fitting fractured from overstress in tension, and each element of the tension-torsion strap pack fractured in overstress. There were no indications of other fracture modes, and no pre-existing cracking, corrosion, or other degradation was observed. The engine air inlet fairing sections also exhibited fractures consistent with overstress. Examination of the blue streaks on the outer surface of the fairing exhibited features consistent with paint transfer from the main rotor blade. For additional information, refer to the complete Materials Lab Factual Report contained in the public docket for this accident.