N26WW
Aerospatiale AS-350BA
The helicopter departed from a helipad located on an offshore oil platform, and was 3 minutes from landing at a refueling helipad on another platform, when the pilot transmitted two distress calls indicating the helicopter was "going down." There were no witnesses to the accident; however, 9 minutes after the distress calls were heard, the helicopter was found floating inverted in 3-4 foot seas. Subsequently, the helicopter sank and was later recovered and examined. Examination of the helicopter revealed no anomalies with the airframe or flight control systems. Examination of the engine revealed that the first and second stage turbine blades were fractured due to extreme heating. One blade of the second stage tubine disk had liberated from its retention slot, and all the blade roots and retention slots of this disk exhibited permanent outboard deformation, due to a combination of centripetal forces from engine operation and excessive heat. In contrast, the blade roots and retention slots of the first stage turbine disk did not exhibit evidence of outboard deformation, most likely since they were located further away from the heat source. The rear bearing assembly (located aft of the second stage disk) was contaminated with coke. The coking suggests that oil was leaking from the engine and migrating from the rear bearing assembly. The aft side of the second stage turbine disk displayed dark stain marks in the form of streaks. A passage exists that would allow oil to flow from the rear bearing to the aft face of the second stage turbine disk. Oil that strikes the disk would flow into the hot stream of gases and auto-ignite, starting a fire. Oil migration can occur if the rear bearing scavenge and vent tubes become blocked; however, the scavenge and vent tubes were checked during the engine examination and were not found blocked.
HISTORY OF FLIGHT On October 28, 2000, at 1637 central daylight time, an Aerospatiale AS-350BA single-engine helicopter, N26WW, was destroyed when it impacted water during a forced landing following a loss of engine power near the East Cameron area, block 200, in the Gulf of Mexico. The instrument rated commercial pilot, sole occupant of the helicopter, received fatal injuries. The helicopter was registered to Textron Financial Corporation Inc., of Fort Worth, Texas, and operated by Tex-Air Helicopters Inc., of Houston, Texas. Visual meteorological conditions prevailed, and a company visual flight rules (VFR) flight plan was filed for the 14 Code of Federal Regulations Part 91 positioning flight. The flight originated from East Cameron 231A (EC231A), an offshore oil platform located in the Gulf of Mexico, at 1630, and was destined for Abbeville, Louisiana, with an intermediate fuel stop at Vermilion 200 (V200), another offshore oil platform. According to dispatch records and the pilot of another helicopter (N118TA), at 1635, the pilot of N26WW stated the he was en route to V200 and was three minutes from landing. At 1636, the pilot of N118TA heard the following distress call: "118, I'm going down." Before the pilot of N118TA could respond he heard a second distress call during which he heard: "118, Whiskey Whiskey going down one mile from you." The pilot of N118TA reported that he could hear a "warning horn," whose sound could not be positively identified, as the pilot of N26WW transmitted the distress call. No other distress calls were received. The pilot of N118TA then initiated a search for N26WW. At 1645, the pilot of N118TA located the accident helicopter floating inverted. According to a global positioning system (GPS) receiver in N118TA, N26WW was located at north 028 degrees 44.28 minutes latitude and west 092 degrees 37.80 minutes longitude, approximately 75 miles southeast of Cameron, Louisiana. The pilot did not observe any survivors or inflated personal floatation devices. Each of the helicopter's six skid floats were observed inflated. One of the floats remained attached to the skid assembly and the remaining five were floating beside the helicopter. At 1650, the float that remained attached to the skid detached, and the helicopter sank. On November 4, 2000, the helicopter was recovered to a secure facility at the Clover Field Airport, Houston, Texas. PERSONNEL INFORMATION On April 26, 1992, the pilot was issued a commercial pilot certificate, with rotorcraft-helicopter and instrument rotorcraft-helicopter ratings. He also held airplane single and multi-engine land ratings. According to a training form, provided by Tex-Air Helicopters Inc., the pilot completed initial training for the AS-350 on June 4, 1999, and underwent his most recent FAR Part 135 competency/proficiency check on July 31, 2000. According to the Pilot/Operator Aircraft Accident Report (NTSB Form 6120.1/2), which was completed by Tex-Air Helicopters Inc., the pilot had accumulated a total of 3,591.4 flight hours, of which 3,481.1 were in rotorcraft, and 856.8 were in the AS-350. On October 6, 2000, the pilot was issued an FAA second class medical certificate, with no limitations. Additionally, the pilot had completed an offshore water survival course on June 19, 1998. AIRCRAFT INFORMATION The 1980-model black, white, and red helicopter (serial number 1229) was equipped with a three bladed, semi-articulated main rotor system and a 681-horsepower Turbomeca Arreil-1B turboshaft engine. The helicopter was operated in Japan from the time of its inaugural flight in 1985 through 1998. On July 23, 1991, the engine (serial number 757) was removed from another AS-350, overhauled, and installed in the accident helicopter. At the time of the overhaul, the engine had accumulated a total of 1,958.06 hours and the airframe had accumulated a total of 3,959.05 flight hours. During the overhaul, the first and second stage turbine discs were replaced and modifications were performed to module three (which houses the high pressure section of the gas generator with the power turbine), which changed its overhaul interval from 2,000 to 2,500 hours. On October 18, 1994, the helicopter was involved in a mid-air collision; which damaged its tailboom and horizontal stabilizer. The records revealed that the helicopter was then stored in Japan for a period of four years. During 1998, the helicopter was imported to the United States. On October 6, 1998, repairs to the damage that was incurred during the mid-air collision were completed, and the helicopter was converted from a B model to a BA model by Arrow Aviation of Broussard, Louisiana. On October 8, 1998, the helicopter was issued an FAA standard airworthiness certificate, at which time it had accumulated a total of 4,779.8 flight hours. The engine had accumulated a total of 2,778.8 hours. On May 18, 2000, the engine's rear bearing oil scavenge pipe was flow checked for clogging. According to the Turbomeca Maintenance Manual (05-10-02), each operator is required to perform this check every 100 hours of engine operation. Clogging of the rear bearing oil scavenge pipe could result in oil residue collecting, in the form of coke, in the area of the rear bearing. Coke is a hard, crystalline residue of turbine engine oil. According to a maintenance log card provided by the operator, the rear bearing oil flow was less than the allowable limits. The engine was removed from the airframe, disassembled to allow servicing of the rear bearing, and subsequently, the engine was reassembled and reinstalled in the airframe. Additionally, the engine was modified to incorporate TU 281, 282, and 283 engine modifications. According to Turbomeca Service Bulletin N 292 72 0215, the modifications "improve cooling of the gas generator rear bearing and reduce the formation of coke deposits in this area." The airframe had accumulated a total of 6,198.0 flight hours and the engine had accumulated a total of 4,197.1 hours when the work was accomplished. On October 10, 2000, the engine underwent its most recent 100-hour inspection, and had accumulated a total of 4,687.4 hours. On October 24, 2000, the airframe underwent its most recent 100-hour inspection and had accumulated a total of 6,731.0 flight hours. On October 27, 2000, 50 gallons of Jet-A fuel were added to the helicopter's 143 gallon capacity fuel tank. According to the operator's daily flight logs for October 27th and 28th, the helicopter had completed 14 flights since the last refueling, and would have had approximately 40 minutes of fuel on board at the time of the accident. At the time of the accident, the airframe had accumulated a total of 6,733.9 flight hours, and the engine had accumulated a total of 4,733.0 hours. According to the maintenance records, all flow checks of the rear bearing subsequent to May 18, 2000, were complied with, and no clogging was found. Additionally, it was determined that module three of the engine had accumulated 2,774.9 hours since overhaul, which was 274.9 hours over its 2,500 hour overhaul interval. METEOROLOGICAL INFORMATION The pilot of N118TA reported the following weather conditions near the accident site; scattered clouds at 5,000 feet, visibility 10 miles in haze, wind from 100 degrees at 12 knots, and seas 3-4 feet. WRECKAGE AND IMPACT INFORMATION A review of an underwater recovery video revealed that the helicopter came to rest on its left side. The rotor brake lever was in the released position, the fuel flow control lever (throttle) was 1/4 out of the flight gate in the emergency operating range, and the emergency fuel shutoff lever was in the open position (normal ready position). The recovered wreckage was examined by the NTSB investigator-in-charge, representatives from Tex-Air, and a representative from American Eurocopter. The cabin area was damaged; however, the instrument panel, collective, cyclic, anti-torque pedals, and rear entry doors remained intact. The nose, windscreen, both front entry doors, and cabin roof were separated from the airframe. All of the belly panels were separated, except for the rear center panel which exhibited hydraulic deformation in the upward direction. The skid landing gear assembly remained attached to the airframe. The engine and transmission remained attached to the airframe; however, the main rotor system separated from the transmission. The tailboom separated from the airframe at the fuselage junction, and was also separated between the horizontal and vertical stabilizers. It was noted that damage to the fuselage corresponded to the artificial horizon's resting position, which was right bank and nose up. The pilot's cyclic and collective controls remained intact; however, deformation of the flight control tubes beneath the cockpit prevented movement of the controls. Continuity was confirmed from the cyclic and collective controls to the transmission deck. The collective's low pitch stop was 19.0mm from the web down position. There is no manufacturer's recommended setting for the low pitch stop; however, three exemplar helicopter's low pitch stops were measured and found to be between 14mm and 16mm. Tail rotor control continuity was confirmed from the cockpit to the tailboom disconnection point, and from the tailboom disconnection point aft to the tail rotor blades. The three main rotor blades remained attached to the main rotor head. One blade (serial number 10533) and a second blade (serial number 10533) displayed trailing edge delamination; however, no leading edge damage was observed. The third blade (serial number 10539) displayed trailing edge delamination, and exhibited a crack at the serial number location that extended longitudinally through the entire blade length. The main rotor system separated from the main gear box at the epicyclic gears. The epicyclic gears remained attached to the main rotor shaft and the sun gear remained attached to the main gear box. The epicyclic gears were corroded and seized. Lubricant was applied to the epicyclic gears and they were rotated by hand. Continuity was established from the epicyclic gears, through the main rotor shaft, to the main rotor blades. The pitch change rods were removed and measured. The yellow (master) was 373.5 mm, the blue was 372.0 mm, and the red was 375.0 mm. The master pitch change rod was within the manufacturer's recommended setting. The main gear box was examined. The magnesium outer casing contained holes that were 4 inches in diameter due to corrosion from salt water immersion. The bevel reduction gear assembly, which was viewed through the holes, was intact, corroded, and seized. The oil pump shaft was intact, and oil was observed in the oil level sight gauge. The tail rotor drive system was examined. The steel shaft (often referred to as the intermediate or short shaft) connecting the engine power shaft to the tailrotor drive shaft separated at the coupling attachment nearest the engine. The engine casing exhibited rotational scoring, consistent with contact with the coupling. The tail rotor blades remained attached to the hub and rotated through the 90-degree gear box when manipulated by hand. The tailrotor gearbox cowling displayed witness marks consistent with contact with a rotating tail rotor drive shaft input coupling. The leading edge surfaces of both tail rotor blades were undamaged; however, one tail rotor blade was fractured near the hub. The engine was externally examined. The compressor displayed no signatures of foreign object damage. The exhaust stack exhibited multiple pea sized dents, which originated from the inside and protruded externally. The engine was sent to the Turbomeca Engine Corporation facility in Grand Prairie, Texas, for further examination. MEDICAL AND PATHOLOGICAL INFORMATION An autopsy was performed by the Lafayette Parish Coroner's Office Lafayette, Louisiana. The cause of death was determined to be "multiple injuries," related to a helicopter crash. An FAA toxicological specimen kit was delivered to the medical examiner for toxicological testing; however, specimens were not received by the FAA in accordance with the directions enclosed in the kit. TESTS AND RESEARCH On November 14, 2000, the engine was examined at the Turbomeca Engine Corporation facility. The axial compressor rotor rotated freely when manipulated by hand. The axial compressor rotor blade tips were curled, and rub marks consistent with blade contact were observed on its shroud. The centrifugal compressor rotor was intact and displayed signatures consistent with corrosion due to immersion in salt water. Rub marks were observed on one half of the blade tips. The combustion chamber was intact. The fuel injectors and igniters were clean, free of debris, and no anomalies were noted with the burn pattern. The first stage turbine nozzle guide vane (NGV) was intact. The first stage turbine disk blades displayed thermal damage (soot covered and charred) and appeared fractured; however, they remained seated in the turbine disk. (Each turbine blade's fir tree fingers (root) are inserted into a retention slot in the turbine disk. The blade is secured to the disk by a blade retention pin. The blade platform is located between the fir tree fingers and the blade airfoil surface.) The second stage NGV was sooted, charred, and displayed metal splatter. The second stage turbine disk was missing one blade. A hole was observed in the NGV, which contained a corresponding witness mark consistent with the size of a turbine blade. A fragment of a blade, which consisted of the blade root (fir tree fingers), was found in the NGV. The blade's outboard portion was not recovered. Additionally, the containment shield (which is fitted around the NGV), in the area of the hole, was melted and deformed. Each of the remaining second stage turbine blades was seated in the turbine disk; however, they displayed thermal damage and appeared fractured. The rear bearing assembly (located just aft of second stage turbine disk) was observed to be contaminated with coke. The rear bearing oil return and ventilation tubes were examined and observed to be free from contamination and debris. The free turbine assembly was intact and rotated freely when manipulated by hand. The reduction gear box was intact, lubricated and displayed no signatures of wear. The rear drive power output shaft (free-wheel shaft), which drives the tail rotor drive shaft, was separated. The accessory gear box was intact and the casing was opened. The accessory gears displayed no wear signatures and the gears rotated when manipulated by hand. The front power output shaft was intact. Each of the engine bearings, from the axial compressor rearward to the output shaft bearing, were examined, and all appeared intact and lubricated. The fuel control unit was placed on a test stand, functionally tested, and no anomalies were noted. Between November 27 and November 30, 2000, the first stage NGV assembly and turbine disk with its attached blades; second stage containment shield, NGV assembly, and turbine disk with its attached blades; one turbine blade that separated from the second stage turbine disk; one bag of debris from the first stage blade path; one bag of debris from the first stage NGV blade path; one bag of debris from the second stage blade path; and the fractured free-wheel shaft assembly were examined at the NTSB Materials Laboratory in Washington, D.C. The first stage NGV, turbine disk, and turbine blades were examined; no material anomalies were noted. The first stage NGV displayed black deposits and thermal damage around its circumference. The blades remained attached to the disk; however, they were fractured between 0.2 and 0.4-inches above the platform. The fracture surface of the blades displayed thermal damage. The blade roots (fir tree fingers) and their respective blade retention slots did not exhibit evidence of deformation. The blade retention pins were secure and not loose. Additionally, the front and aft surfaces of the turbine disk were clean and unstained. The second stage containm
the loss of engine power due to an internal engine oil leak that started an internal engine fire and the pilot's inadequate autorotation which resulted in a hard landing. A contributing factor to the accident was the rough water condition.
Source: NTSB Aviation Accident Database
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