NTSB Factual Report

On June 8, 2006, about 1232 central daylight time, a Fairchild Hiller FH-1100, N4035G, registered to and operated by Helicopters of NW Florida, Inc., experienced separation of a tail rotor blade during cruise flight near Navarre, Florida. Visual meteorological conditions prevailed at the time and no flight plan was filed for the 14 CFR Part 91 personal flight from Pullum Pad, Navarre, Florida, to a business located in Century, Florida. The helicopter was substantially damaged and there were no injuries to the commercial-rated pilot or pilot-rated passenger. The flight originated about 1225, from Pullum Pad, Navarre, Florida.

The pilot stated that during level flight at 600 feet and 100-110 miles-per-hour, flying westbound along the intercoastal waterway, he heard a bang and "felt a slight shudder of the aircraft." He immediately entered a descent for a precautionary landing to a beach, and while descending, did not experience any loss of directional control and no abnormal vibration. He maintained the power setting at 100 percent, and when the flight was at 50 feet, he entered a "gentle" flare and brought the helicopter to level flight at 30 feet. At that time, the helicopter began to yaw to the right which left anti-torque pedal input did not correct. He then immediately retarded the throttle, and executed a hover autorotation to shallow water along the beach. The helicopter was removed from the water and placed on the beach for further examination.

Examination of the helicopter by an FAA airworthiness inspector revealed one of the tail rotor blades was separated and not recovered, while the other tail rotor blade assembly remained attached to the hub. Approximately 6.5 inches of the main spar of the ventral fin was separated. The separated section of ventral fin structure remained secured to the tail rotor gearbox, which remained partially secured to the helicopter by the anti-torque cables. There was no evidence of bird contact on any remaining portion of the helicopter or tail rotor blade. The tail rotor hub assembly and gearbox were retained for further examination.

Examination of the tail rotor gearbox and tail rotor hub assembly was performed by the FAA, and attended by a representative of the helicopter manufacturer. Visual examination of the tail rotor gearbox revealed the output shaft and pitch change rod were bent 15 degrees. The input shaft flex coupling was broken; rotational scoring was noted on the input shaft flange. Disassembly of the gearbox revealed no impact signatures on the helical gear faces. The unit was properly serviced, and no signs of contamination were noted. No abnormal wear was noted. Examination of the tail rotor hub assembly pertaining to the separated tail rotor blade revealed the laminated stainless steel plates of the tension-torsion bar were fractured; 5 semi-circular pieces of the stainless steel laminated plates were recovered. The fractured tension-torsion bar assembly was retained for further examination. The remaining tail rotor blade assembly was removed from the tail rotor hub, and the laminated stainless steel plates that comprise the tension-torsion bar assembly were not fractured or failed. Visual examination of the holes on each end of the laminated stainless steel plates revealed "manufacturing (stamping) marks on the individual straps." The intact tension-torsion bar assembly was also retained for further examination.

Examination of the fractured and intact tension-torsion bar assemblies was performed by the NTSB Materials Laboratory located in Washington, D.C. No markings (serial number) was noted on either tension-torsion bar assembly. Examination of the fractured tension-torsion bar assembly revealed only 6 of the 11 laminated stainless steel plates remained secured to the bolt at the tail rotor hub end. The opposite ends of the all 6 plates were fractured near the eye. Examination of the fracture surfaces of the 6 stainless steel laminated plates and also of the 5 fractured semi-circular pieces revealed none of the fracture surfaces matched, i.e., the received pieces from the fractured tension-torsion bar constituted samples from all 11 stainless steel laminated plates from the fractured tension-torsion bar assembly. Further examination of the fracture surfaces of the 11 laminated plates revealed 8 of the 11 plates exhibited fatigue progression. Fatigue was noted in 1 of the 11 plates on both ends of the plate. The fatigue initiated on the inner diameter surface of the eye and progressed outward. "...normal to the long axis of the strap." The extent of the fatigue ranged from an estimated 10 percent of the fracture surface to about 90 percent. The remainder of the fracture surface was ductile dimpled overstress, generally oriented on a slant plane through the strap thickness. The inner diameter surfaces of the straps mostly showed circumferential marks consistent with reamed surfaces; however, all of the fatigue origin areas were located in "...rougher textured regions consistent with sheared surfaces from original manufacture of the straps...." The inner diameter surfaces of the intact eyes of both stainless steel laminated plates exhibited "...similar (reamed and sheared) surfaces on the intact eyes of the straps. Additionally two small cracks were uncovered in separate straps from the fractured TT bar. No cracks were visually apparent in the eyes of the intact TT bar." The stainless steel plate material was consistent with AISI type 300 series stainless steel.

The helicopter was manufactured on October 18, 1982, and had accumulated 374.0 hours since manufacture at the time of the accident. The helicopter was involved in a previous accident on February 18, 1985, which was investigated by the NTSB and assigned case number DEN85LA079. The helicopter was not operated from February 18, 1985, to February 5, 2003. The helicopter was operated 4.3 hours between February 5, 2003, and December 23, 2004, and was operated for 164 hours between December 23, 2004, and the date of the accident. The maintenance records indicated that the tension-torsion bar assemblies serial number were 2019, and 2059, and the part number for each was 24-55106. There was no record that the tension-torsion bar assemblies were replaced, or repaired since manufacture.

On August 5, 1977, the FAA issued Airworthiness Directive (AD) 77-07-08, which indicated inspection of tension-torsion bar assemblies serial number prior to 1922; the AD was not applicable to the accident tension-torsion plates based on the serial number of the plates. The AD called to inspect the inner diameter of the holes for evidence of improper surface finish, in accordance with the manufacturer's Service Bulletin (SB) SBFH1100-55-2A, dated January 19, 1977.

Review of service bulletin SBFH1100-55-2A, revealed that unacceptable findings from the inspection include the inner surface of the hole is broken with transverse tool marks, and pitting or damage that breaks through one edge of the hole. As previously reported, the fatigue origin areas of the accident tension-torsion bar assemblies were located in rougher textured regions consistent with sheared surfaces from the original manufacture of the straps.

The NTSB retained torsion bar assemblies for both tail rotor blades were released to Georges Van Nevel, president of FH 1100 Manufacturing Corporation, on October 10, 2007.

NTSB Probable Cause

The inadequate manufacturing of the plates of the tail rotor tension-torsion bar assembly resulting in fatigue failure.