On April 20, 2005, about 0745 eastern daylight time, a Bell 206B, N2285B, registered to and operated by Heliworks, Inc., rolled over while lifting off from the Everglades near Coral Springs, Florida. Visual meteorological conditions prevailed at the time and no flight plan was filed for the 14 CFR Part 135 local, other work use flight from Fort Lauderdale Executive Airport, Fort Lauderdale, Florida. The helicopter was substantially damaged and the commercial-rated pilot and two passengers were not injured. One passenger sustained serious injuries. The flight originated about 0700, from the Fort Lauderdale Executive Airport. Use your browsers 'back' function to return to synopsisReturn to Query Page
The pilot verbally stated that after takeoff, the flight proceeded approximately 15 miles west to Sawgrass Park where he landed and picked up three workers. He completed a load manifest and computed the weight and balance. He then proceeded to a site for vegetation eradication, and after landing, the workers got out, sprayed, then returned. He then departed again to another site where one of the workers got out, sprayed, and returned to the helicopter. He lifted up straight and the right side "popped up fast." He lowered collective and applied right cyclic to correct the roll which had no affect; the helicopter rolled onto its left side. He further reported he did not perceive a problem with the helicopter or flight controls.
Examination of pictures provided by the operator revealed the helicopter was resting on its left side partially submerged. The tailboom was fractured but in close proximity to the wreckage and damage was noted to the bottom of the fuselage just aft of the aft crosstube. One of the main rotor blades was visible. According to FAA personnel, during recovery, the helicopter was dropped from a height of approximately 20-30 feet.
National Transportation Safety Board examination of the helicopter following recovery revealed the main rotor mast was fractured just below the static stop contact zone; the fracture surface circumferentially exhibited 45-degree shear lips. Both main rotor blades were fractured; 45 degree shear lips were noted on the fracture surfaces of both blades. One of the two main rotor blades was fractured approximately 152.5 inches from the centerline of the attach bolt; blue colored paint was noted on the leading edge of the blade. The other blade was fractured approximately 148 inches from the centerline of the attach bolt; blue colored paint was noted on the upper surface of the blade. The tailboom was separated at approximately boom station 63. One section of tailrotor drive shaft was displaced due to aft displacement of the 2nd tailrotor drive shaft bearing.
Examination of the left rear seat revealed the seatback cushion was not in-place, and the shoulder harness was connected to the lapbelt, but the male and female ends of the lapbelt were not connected. Examination of the right rear seat revealed the shoulder harness was connected to the lapbelt but the male and female ends of the lapbelt were not connected.
Examination of the collective flight control system revealed control tube assembly continuity from the cockpit to the lever assembly; a fracture was noted to the bellcrank P/N 206-001-568-001, near the area where the tube assembly connects. No evidence of preimpact failure or malfunction was noted on the fracture surface of the bellcrank assembly. Examination of the cyclic flight control system revealed control tube assembly continuity from the cockpit to each bellcrank assembly. Each control tube assembly was fractured between the bellcrank assembly and the inner ring assembly. Both fractured control tubes were bent and exhibited "D" shaped deformation in the area of the fracture surface. One pitch link assembly was fractured between the attach point on the outer ring assembly and the attach point near the main rotor blade. The other pitch link assembly remained connected to the attach point near the main rotor blade, but the other end was not connected to the outer ring assembly. The end of the pitch link that was separated from the outer ring assembly still had the securing hardware and bearing connected to the end of the link. Examination of the outer ring assembly revealed one of the pitch link assembly attach point fitting was fractured; no evidence of preimpact failure or malfunction was noted to the fracture surface. The left and right cyclic, and the collective servo actuators were removed from the airplane for further examination at the helicopter manufacturer's facility with FAA oversight.
Bench testing of the left cyclic servo actuator (S/N 6608) revealed the relief valve pressure "cracked" at 810 psig (specification is 825 to 895 psig test port pressure). The relief valve closed at 570 psig (specification is that it must close within 120 psig of the cracking pressure). During the "Manual Operation Test", the manual force to move the cylinder was 44 pounds (specification is 26 pounds or less). All other sections of the test procedure were within normal limits. Testing of a sample of fluid revealed the particle count was greater than specified for all channels. Bench testing of the right cyclic servo actuator (S/N 2310) revealed with respect to the "Servo Valve Leakage Test", the leakage amount was 50cc/minute (specification is 20 cc/minute). All other sections of the test procedure were within limits. Testing of a sample of fluid revealed the particle count was greater than specified for all but one of the five channels.
Bench testing of the collective servo actuator (S/N 6576), revealed that with respect to the "Manual Operation Test", the unit did not move when a force of 40 pounds was applied (specification is 26 pounds or less). A force of 600 psi (normal 206B system pressure) was applied to the pressure port with the return port open and the cylinder would cycle. A "thick black material" was noted extruding from each end of the barrel. Further testing of the collective servo actuator with respect to the "Un-Boosted Force Test" revealed the peak force to cause movement of the piston ranged from 134 to 92 pounds in the retract direction and 88 to 121 pounds in the extend direction. The servo was then disassembled which revealed a localized area of "...fresh burnishing" of one side of the inboard gland bore. The piston rod was checked for straightness using "V-blocks" and a dial indicator and the total dial indicator run-out was .0065 inch.
The helicopter minus the retained components was released to David E. Gourgues, Regional Manager for CTC Services Aviation (LAD) Inc., on November 10, 2005. All NTSB retained components were also released to David Gourgues, on March 29, 2006.