Lancaster, CA, USA
N177SJ
San Joaquin Helicopters OH-58A+
Following a loss of engine power, the helicopter impacted the ground during the application of chemicals to an agricultural field. Post accident examination of the engine showed that the power turbine to pinion gear coupling had fractured. Metallurgical examination of the affected components showed that the coupling failed in fatigue, which effectively separated the engine from the main rotor gearbox. The fatigue was the result of a misaligned power turbine outer shaft, where the axis of the turbine shaft was angled slightly relative to the coupling and pinion axes. The turbine section had been removed from another engine due to metal contamination. The turbine section was repaired and subsequently installed on the accident helicopter on August 11, 2004, where it operated for 770.3 hours prior to the accident. Weight and balance calculations showed that the helicopter was in excess of its certificated maximum gross weight by 174 pounds.
HISTORY OF FLIGHT On August 10, 2005, at 0600 Pacific daylight time, a San Joaquin Helicopters OH-58A+ restricted category helicopter, N177SJ, impacted terrain while performing aerial pesticide application to onion crops in Lancaster, California. San Joaquin Helicopters, who was also the registered owner of the helicopter, was operating it under the provisions of 14 CFR Part 137. The commercial pilot sustained fatal injuries; the helicopter was destroyed. Visual meteorological conditions prevailed and no flight plan had been filed. The pilot departed about 3 minutes prior to the accident and was completing the first pesticide application of the day when the accident occurred. According to a local farmer who was assisting in the ground portion of the aerial application, the pilot departed from a fuel and chemical loading truck about 1/2-mile south of the accident site. He estimated that the pilot flew the helicopter about 60 to 65 miles per hour and 20 feet above ground level. He maneuvered the helicopter upward to clear a house surrounded by 40-foot tall trees. The helicopter then started downward toward the field and went out of the view of the witness. The witness then drove around a corner and saw the helicopter continue downward toward the field until the nose of the helicopter impacted the ground. Upon impact, the helicopter tumbled down the field before coming to rest. As the helicopter descended toward the field, the witness did not hear any unusual noises coming from the engine, and the approach path to the field appeared normal. PERSONNEL INFORMATION The pilot held a commercial pilot certificate for single engine airplanes and helicopters. A second-class medical was issued in March of 2005, and held the restriction that the pilot must wear corrective lenses. The medical was issued on the condition of a waiver for color blindness. The operator reported that the pilot had a total flight time of 4,089 hours in helicopters with 50 hours in the last 90 days. A copy of the pilot's personal flight logbook was obtained from his family and recorded a total of 4,481.2 flight hours in helicopters. The pilot entered the last 6 entries by monthly flight hour totals, with the last entry dated July 2005. AIRCRAFT INFORMATION The helicopter was maintained in accordance with the Federal Aviation Administrations (FAA) approved San Joaquin Helicopters instructions for continued airworthiness report (Report SJH 97-001). The last annual inspection was completed on March 23, 2005, at a total airframe time of 4,894.3 hours. The last inspection (for continued airworthiness) was completed on June 22, 2005, at a total airframe time of 5,211.00 hours, and a total engine time of 2,446.9 hours. The helicopter had accrued 99 hours since the last inspection. At the time of the accident, the engine had approximately 2,540.8 hours. On August 31, 2004, the turbine was installed on the engine at a total engine time of 1,770.5 hours. The turbine section had been removed from another engine for "metal contamination." The turbine was inspected and subsequently installed on the helicopter where it operated for 770.3 hours prior to the accident. San Joaquin Helicopters modified the accident helicopter from a Bell OH-58A to a Bell OH-58A+, and a new type certificate was issued on February 4, 1998. The modification from the original type design included the installation of the four following items: 1. Lead-acid battery installation 2. Ag-Air systems installation 3. ISOLAIR Model 3900-OH58 spray system installation 4. Flight hour recording meter The helicopter was restricted to agricultural operations in accordance with 14 CFR Part 21.25. WRECKAGE AND IMPACT INFORMATION The National Transportation Safety Board investigator-in-charge (IIC), the FAA accident coordinator, and representatives from Bell Helicopter Textron, and Rolls Royce, all parties to the investigation, responded to the accident site on August 11, 2005. The main wreckage came to rest at these approximate global positioning coordinates: N 34 degrees, 42.437 minutes W 118 degrees, 01.952 minutes, at an elevation of about 2,400 mean sea level (msl). The helicopter impacted a field of onion crops that ran from south to north. The wreckage area covered approximately 80 yards, spanned a width of 40 feet, and the debris field was in a northerly direction. On the southeast corner of the property was a house surrounded by 40-foot tall trees. The crops began about 200 feet from the house. Powerlines were located on the eastern edge of the field. The first identified point of contact (FIPC) was one of the nozzles from the chemical spray system, located about 75 feet from the start of the field's southern side. North from FIPC was the toe portion of the right skid tube. The toe of the skid was buried approximately 2 feet in the soft agricultural soil, and positioned vertically from the ground. Forward and to the right of the skid was a rectangular imprint, similar in size and shape to a main rotor blade. The top portions of the onion stalks in the area of the rectangular imprint were cut at 45-degree angles. Branching outward and to the right from this area were additional rectangular imprints. The main rotor blades and hub assembly were just forward of the imprints and about mid-debris field. One of the tail rotor blades, and the right anti-torque pedal were located 50 feet to the left of the perpendicular imprints. The crops between the main rotor assembly and the tail rotor blades were noticeably disrupted and contained various helicopter debris leading up to the main wreckage. Scattered throughout the debris path were the left door, Plexiglass shards, chemical spray nozzles, the seat back, and cockpit paperwork. The main structure of the helicopter was resting on its right side along a northerly heading. The cockpit structure and instrument panel were crushed and separated from the cabin portion and just forward of the main structure. The cabin area remained intact. Forward 30 feet and 15 feet to the left of the main structure, was the center section of the tailboom that was separated from the main structure one foot aft on the tailboom. Approximately 20 feet forward and 10 feet right of the main structure, was the tail rotor assembly still connected to the aft portion of the tailboom. Portions of the chemical spray system on the helicopter were found throughout the debris field. The majority of the system had separated from the fuselage, as well as from the spray system attachment frame. MEDICAL INFORMATION The Los Angeles County Coroner completed an autopsy on the pilot. The FAA Bioaeronautical Research Laboratory completed toxicological testing on specimens of the pilot. The results were negative for all tested drugs, volatiles, cyanide, and carbon monoxide. TESTS AND RESEARCH The recovered wreckage was examined at Aircraft Recovery Services, Inc., on August 12, 2005, by the NTSB investigator and representatives from Bell Helicopter Textron and Rolls Royce. Airframe The cockpit structure of the helicopter was crushed and torn away from the remaining cabin structure during the accident sequence. The right side appeared more damaged than the left side. The left seat pan remained intact and the cockpit floor containing the anti-torque pedals was torn away. The right seat (pilot's seat) pan and cabin structure were separated from the seat frame. Three of the four pilot's seat belt attachments were fractured. On the left side of the seat belt, both attach points were fractured. The top right attach point was fractured, but the lower right attach point was intact. The right shoulder harness webbing had separated. The pilot's restraint system was removed from the helicopter and sent to the NTSB Materials laboratory for further examination. The flight controls, or portions of them, were traced throughout the wreckage. According to the helicopter manufacturer's representative, all fractures were consistent with overload. The tail rotor driveshaft was torsionally fractured at the sixth segment of the shaft. The tail rotor pitch change rod was manually actuated at the separated tail rotor section and produced movement to the tail rotor assembly. All hanger bearings and fractured tail rotor driveshaft sections rotated freely, and all Thomas couplings were intact with no missing hardware. The oil was drained from the tail rotor gearbox and did not contain noticeable contaminants. The tail rotor gearbox chip detector was removed and no chips or debris were noted. The main rotor blades were twisted and deformed and the swashplate assembly was connected to the main rotor hub through one pitch change link. Drive system continuity was established throughout all sections of the drive train. Both chip detectors were removed and no debris or chips were observed. Engine The Allison T63-A720 (Commercial Version 250-C20C, Serial Number CAE406008) four-stage turbine engine was examined. The combustion section was crushed on its lower left side. The N1 and N2 rotor systems would not rotate. Pieces of the crops and soil were found in the engine inlet area. The engine was rated at 420 shaft horsepower. The upper and lower chip detectors were removed and did not contain pieces of metal debris. Five ounces of oil were drained from the accessory gearbox. The oil was dark in color. The fuel pump filter element was removed and found free from debris. The oil filter element was removed from the accessory gear housing and contained deposits similar in color to carbon. Fuel was drained from the fuel line connecting to the fuel nozzle to the fuel control unit. The fuel was clean and free from debris. The engine was disassembled on September 15, 2005, at the Rolls-Royce facility in Oakland, California, with an NTSB investigator and a representative from Rolls Royce in attendance. Following the removal of the external components, the turbine was separated from the gearbox. Investigators noted that the power turbine outer shaft remained intact. The power turbine to pinion gear coupling was fractured and separated at the turbine end. The coupling displayed the following markings: Extex E6870832K, SN ACR3052, FAA PMA (Part Manufacturing Approval). The number 5 bearing cage was distorted and the number 5 bearing race was cracked. Technicians observed no metal splatter in the turbine area. A technician inserted a 1/4-inch drive into the tachometer drives and the N1 and N2 gear trains rotated. The compressor was removed from the gearbox and the compressor assembly rotated freely. The spur adapter gear shaft appeared intact with no damage. The splines on the fuel control, power turbine, governor, and engine driven fuel pump rotated freely. The compressor case halves were removed and the vanes were intact. Slight rub marks were present on the fifth stage compressor case halves. At the completion of the disassembly, no mechanical anomalies were noted. The fuel control unit was examined at Honeywell under the auspices of a Federal Aviation Administration inspector on November 1, 2005. Representatives from Rolls Royce and Honeywell, parties to the investigation, were also present. Due to impact damage, no speed could be input; however, the unit reacted normally to pressure inputs. The power turbine governor was also examined. Functional testing found no condition that would have prevented normal operation. The power turbine to pinion gear coupling, the power turbine outer shaft, and the number 5 bearing were sent to the NTSB Materials Laboratory for further examination. Metallurgical Examinations Engine Component Examination The NTSB Materials Laboratory performed an examination of the fractured power turbine to pinion gear coupling, power turbine outer shaft, and the number 5 bearing (Report No. 06-074). A copy of the full report can be found in the docket material for this accident. The examined components were located, by design, between the engine's accessory gearbox section and its turbine section. The forward end of the coupling is normally connected to the pinion gear by a spline, and the rear end of the coupling is connected to the turbine shaft by a similar spline. The coupling was fractured in two pieces with the rear portion being smaller than the forward portion. The turbine shaft is connected to the turbine and is supported by the number 5 bearing. The direction of rotation for the turbine shaft and coupling is clockwise as viewed looking forward. In summary, the report indicated that the forward portion of the coupling displayed light indications of wear on the contact face of the splines. A circumferential inward bulge was observed and measured. The smallest bulge diameter was below engineering specifications and the area immediately forward of the bulge exceeded engineering diameter specifications. The rear section of the coupling showed a circumferential outward bulge adjacent to the fracture face and mechanical damage to the splines. The diameter of the bulge exceeded engineering specifications. The surface of the coupling had been treated with a black oxide finish. On the rear coupling section, circumferentially oriented bands where the black oxide finish had been rubbed off were present. The width of these bands varied around the circumference and was present on one side of the coupling but not the other side. Mechanical damage to the splines was present to the rearmost portion of the splines. According to the metallurgist, lines on the fracture face were consistent with fatigue crack propagation. The turbine shaft was examined. The exterior surface displayed heat discoloration where the number 5 bearing and spacer were normally located. To the rear of the discoloration, a dark deposit was observed on the surface. The forward face of the turbine shaft, is normally in contact with the pinion. The forward face displayed a ring of material deformation to the aft from approximately mid thickness to the outer surface, with a corresponding outward deformation of the outer surface. The exterior surface of the shaft also displayed the disturbed surface. The inner surfaces of the shaft displayed rubbing, heat, and deformation features and a circumferentially oriented material deposit. The internal splines were discolored and deformed. The deformation on the splines was located on both flanks, with the greatest deformation located at the rear end. The number 5 bearing was examined. The outer race was fractured and a secondary fracture had propagated rearward. Markings consistent with an overstress fracture were observed. Examination of the outer surface of the outer face revealed no distinguishing marks. Examination of the inner surface revealed rub marks on the flat surface forward of the ball groove. The disassembly of the bearing revealed flakes of white metal in the ball grooves and wrapped around some of the balls. The outer surface of the bearing cage displayed circumferential scoring adjacent to its forward and rear faces. The scoring near the forward face was more severe than that at the rear. Blue discoloration at the rear of the forward scored surface was noted. Examination of the forward face and inner surface of the bearing cage also revealed indications of overheating. Pilot Restraint System Examination The pilot's seat belt and shoulder harness assembly were submitted to the NTSB Materials Laboratory Division for examination. Stenciled characters, 1/2-inch high, were present on the left seat belt, adjacent to the latch fitting. The letters "INST" and "SEPT" and the numbers "77" were identified. The examination showed that the right shoulder portion of the harness had separated, approximately 3 inches from the stitching that joined the right portion with the left portion. The inner surface of the separation displayed a darker coloration than the outer surface. The webbing of the harness was woven of strands, each consisting of numerous filaments that had been fused together to form a hard lump, features normally considered as consistent with an overload event. In an overload event, the elongation deformation of t
The pinion to turbine shaft coupling failed in fatigue due to a misaligned turbine shaft, which resulted in a loss of engine power.
Source: NTSB Aviation Accident Database
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