CEN12FA001

The in-flight separation of the main rotor mast following a mast bumping event. The reason for the mast bumping event could not be determined due to postaccident damage.

The helicopter was on a cross-country flight when it experienced a separation of the main rotor mast 8 inches below the teeter bolt, and the main rotor blade assembly separated from the flying helicopter. Subsequently, the helicopter impacted terrain and a postimpact fire ensued. Examination of the mast revealed fracture features consistent with overload failure and mechanical damage indicative of mast bumping. An examination of the airframe, engine, and remaining systems revealed no preimpact anomalies. The reason for the mast bumping event could not be determined due to the amount of thermal damage to the wreckage.

The Clinical Laboratory of the Black Hills, located in Rapid City, South Dakota, performed an autopsy on the pilot on October 3, 2011. The autopsy concluded that the cause of death was due to "blunt trauma injuries" and the report listed the specific injuries. The FAA, Toxicology Accident Research Laboratory, located in Oklahoma City, Oklahoma, conducted toxicological testing on the specimens that were submitted. Carbon monoxide and cyanide tests were not performed. Tests for ethanol were negative. Losartan, a blood pressure medication, was detected in the liver. The main rotor mast, hub, and gear box assemblies were sent to the NTSB Material Laboratory in Washington, DC, for further examination. The inboard end of each main rotor blade was attached to the hub. The enclosure for the gear box assembly and the mast were made from non-ferrous metal. The enclosure for the gear box and other nonferrous pieces in the general area of the gear box showed sagging deformation features consistent with heat damage from exposure to fire. The wall portion of the gear box in certain areas exhibited globule-like features consistent with re-solidified metal. The mast in the area of the fracture exhibited severe bending deformation. The cover for the mast in the area between the hub and gear box also showed evidence of bending deformation and contained a longitudinal fracture in the wall. Portable binocular microscope examination of the ferrous pieces such as the main structural tubes, attachment points for the control surfaces (clevis bolts) and non-ferrous pieces revealed the fracture faces were on slant plane relative to their longitudinal axes and contained dimpled features consistent with overstress separation. The noted signatures were consistent with a mast bump event. The Rolls-Royce RR300 Engine Monitoring Unit (EMU) was examined at the NTSB vehicle recorder laboratory. The EMU records the engine's gas temperature, torque, N1, and N2 speeds. Due to the extreme heat and fire damage, no data was recovered from the device. A teardown examination of the engine was conducted, under the oversight of the NTSB investigator-in-charge (IIC), at Roll-Royce Corporation in Indianapolis, Indiana. The examination revealed the compressor backplate assembly stationary seals, and the impeller rotating seals, exhibited 360 degree rotational scoring. In addition, the P/T-4 nozzle exhibited localized rub area on both the P/T-3 and P/T-4 blade tracks. The inspection of the engine did not reveal any pre impact failures or conditions that would prevent the engine from normal operation. The engine's N1 and N2 coupling adapters were sent to the NTSB Materials Laboratory for further examination. The N1 coupling adapter contained an internal spline at each end. A circumferential fracture split the part into two pieces. The smaller piece contained a longitudinal fracture. The circumferential and longitudinal fractures intersected the internal spline portion at one end of the shaft. The N2 coupling adapter contained only a circumferential fracture. The pieces were ultrasonic cleaned and examination of the fracture faces revealed that they were on a slant plate relative to their longitudinal axes and contained dimpled features consistent with overstress separation. On February 21, 2012, Robinson Helicopter Company released, R66 Service Bulletin SB-03. This service bulletin addressed the possibility of machining burrs being present in certain helicopter hydraulic servos. NTSB investigators removed and examined the three hydraulic control servos from the accident helicopter at Robinson Helicopter Company on November 1, 2012. The examination did not reveal any preimpact anomalies. Multiple swab samples were taken from the wreckage and sent to the Smithsonian National Museum of Natural History for DNA and Microscopic analysis. All samples were negative for bird remains. At 1255, the automated weather observing system at PHP, located 3 nautical miles northeast from the site of the accident, reported wind from 150 degrees at 6 knots, 10 miles visibility, clear of clouds, temperature 84 degrees Fahrenheit, dew point 39 degrees Fahrenheit, and a barometric pressure setting of 30.00 inches of Mercury. The helicopter wreckage and debris came to rest on rolling ranch land and was spread out over an area approximately 1,520 feet long by 600 feet wide. The main rotor head, with attached blades, came to rest 513 feet from the main wreckage. The main wreckage consisted of the fuselage, engine, and tail rotor assembly. The airframe was severely damaged by fire and impact forces. A few pieces of airframe and numerous pieces of Plexiglas were located away from the main wreckage and were not fire damaged. Two portions of the right forward door frame exhibited scoring and deformation similar to the shape of the main rotor blade leading edge. The right front seat bottom had a lateral cut/impact mark near the forward edge and was detached from the helicopter. The controls and tunnel structure exhibited what appeared to be a cut/gash in a vertical line, from the right side, approximately 5 inches forward of the collective pivot. A section of the door frame from the left door exhibited large curved dents, one with paint smearing, on the inside surfaces. The cabin area was nearly consumed by a post-crash fire. All removable controls were found installed. The engine fuel control was found in the full "ON" position. The collective was in a full up position, based on the friction slider. The airframe fuel cutoff valve was extended ¼ inch from its full down position. The remains of what appeared to be a steel shot bag were found beneath the cabin remains. The forward 1.5 bays of tailcone were consumed by fire. The remainder of the tailcone was intact aftward to the last bay, which was separated at the forward end. The empennage had separated from the tail cone and sustained light fire damaged. The small horizontal stabilizer at the base of vertical stabilizer was damaged on left side. The tail rotor visual guard separated from aircraft and fractured in 4 places. Both skid toes and both forward struts separated from the skids. Both rear struts remained attached to skids. The forward cross tube was straight in the horizontal plane and bent forward at the left end in the longitudinal plane and burned through at left side. The rear cross tube destroyed was by fire. The left rear strut exhibited a deep dent on the upper end inboard side. The fuel system was mostly destroyed by fire. The fuel vent tubes had been consumed and only one rollover vent fitting could be located. Investigators were able to blow air through this vent in the upright and inverted positions. The fuel bladder was consumed by fire. The finger screen was recovered and was clear of any debris. The D205-21 fuel hose was found intact, connected to the F670-1 valve. All rod ends for the main rotor flight controls were accounted for and secure to their mounting device, other than at components that were destroyed by fire. The disconnects in the push/pull tubes were consistent with overload and or thermal damage. The servo to swashplate push/pull tubes were bowed. All rod ends for the tail rotor flight controls were accounted for and secure to their mounting device, other than at bellcranks that were destroyed by fire. The disconnects in the push/pull tubes were consistent with overload and or thermal damage. The sprag clutch was subjected to fire and did not rotate. The engine output flex coupling was intact but bent. The F642-1 engine to main rotor gearbox driveshaft was separated at both ends at the yoke weld. The main rotor gearbox input coupling was intact. The main rotor gearbox exhibited thermal damage and had two holes in the case. The output shaft could only be rotated a few degrees. The main rotor gearbox oil filter was not located. In addition, the gearbox rubber mounts were consumed by fire. The main rot