New York, NY, USA
N453AE
EUROCOPTER EC 130 B4
The helicopter was being operated on a revenue sightseeing flight when the accident occurred. Approximately 8 minutes into the flight, about 350-400 feet above the Hudson River while on an approach to land, a 20-inch section of the composite main rotor blade trailing portion, aft of the spar, fractured from the rest of the blade. The pilot reported an immediate decay in main rotor speed with a prominent and abnormal vibration. She also saw a piece of debris, most likely the liberated piece from the main rotor blade, fly from the left rear of the helicopter forward, past the cabin. She made an emergency autorotation onto the water after activating the emergency float system. The helicopter landed upright on its floats; however, the main rotor blades struck the water and the tail boom, resulting in substantial damage to the tail boom. The occupants were rescued by boaters and were not injured. Detailed examination of the helicopter revealed no evidence of flight control system component failures or malfunctions other than the main rotor blade fracture. The main rotor blades (part number 355A-11-0030) were manufactured from glass fiber reinforced composite material with a foam core. From the leading edge to trailing edge of the blade, the blade is constructed with a spar, wedge-shaped foam core, trailing edge roving, and trailing edge tab sandwiched between skin layers and skin reinforcement layers. The skin, skin reinforcement, and trailing edge tab layers are made with glass fabric reinforcement. The trailing edge roving is made of unidirectional glass fibers aligned approximately parallel with the spanwise direction of the blade. The main rotor blade that fractured was examined at the manufacturer’s facility with oversight by France's Bureau of Investigation and Analysis (BEA), and results of the examination were reviewed by the NTSB Materials Laboratory. The National Transportation Safety Board Materials Laboratory also examined sectioned pieces of the blue and yellow main rotor blades submitted by Eurocopter. Physical and microscopic examination of the main rotor blade showed that the fracture was due to fatigue cracks that initiated near the trailing edge of the blade near blade station 1300. It was discovered that the fatigue cracking most likely occurred due to out-of-specification deviations in the alignment of the trailing edge roving fibers within a transition region where the trailing edge roving shifts toward the trailing edge and where skin reinforcement layers end. In the areas of the deviations, the unidirectional fibers of the roving were not properly aligned with the spanwise direction of the blade, likely resulting in localized changes in stiffness at the trailing edge. With this type of fiber misalignment, some of the longitudinal stresses normally carried by the roving layers would be shed to adjacent skin and trailing edge tab areas, which can result in fatigue cracking in these adjacent layers. The undamaged fracture features in the trailing edge roving of the blue main rotor blade section revealed fiber fractures with mirror fracture surfaces across the fiber diameters indicating they were substantially weakened when they fractured. The mirror fractures in the fibers could be evidence of progressive fracture through the trailing edge roving due to fatigue or environmental attack. The fracture surface overall was relatively rough and did not form a flat plane typical of fatigue fracture in tension, such as observed in some areas of the skin. However, some of the fibers had a step at one side or both sides of the fracture that could suggest a mixed mode of loading including tension and transverse shear, which could theoretically explain the overall roughness of the fractures. Also, as cracking progresses the matrix material surrounding the fibers will crack, allowing environmental exposure that could potentially weaken the fibers. The fractographic evidence indicates the trailing edge roving was likely significantly weaker than expected. The weakening of the trailing edge roving fibers could occur due to mechanical damage to the surfaces of the fiber during the fatigue cracking process or could be due to an environmental attack of the fibers as cracking progressed. In either case, the strength of individual fibers was less than expected in a large area of the trailing edge roving. The total extent to which the trailing edge roving was weakened is unknown due to post-fracture damage and the unknown extent to which the trailing edge roving might have cracked before final fracture. In previous cases of skin cracking, an intact trailing edge roving was required to maintain crack stability. In this case however, the evidence shows the trailing edge roving was likely significantly weaker than expected. The previous cases cannot be used as evidence for crack stability in this case due to the weakened condition of the trailing edge roving. Given the extent of cracking in the skin and evidence of weakening in the trailing edge roving, it is likely that the cracking at the trailing edge of the blue blade had proceeded to an extent to cause fracture of the blade. The deviation in the trailing edge roving fibers occurred during the manufacture of the accident blade and it is likely that the trailing edge fibers shifted during the curing process. The skin layers of the blades are somewhat translucent after curing, and manufacturing records showed that the inspector who performed the visual examination after curing flagged the transition area of this blade for a radiographic inspection. It is possible that the inspector observed an anomaly in the trailing edge roving; however, the out-of-specification deviations in the trailing edge roving were not detected by the radiographic examination. A record search by the blade manufacturer of 9,761 similar blades revealed that one other blade had been flagged for radiographic inspection near blade station 1300 during visual examination and subsequently passed radiographic inspection. That blade was returned to the manufacturer’s facility and examined. The blade was sectioned, and the trailing edge roving showed no significant deviations from the as-designed position. The two other blades from the accident helicopter were also examined at the manufacturer. The trailing edge roving in these blades did not show any significant deviations from the as-designed position. The main rotor blade was rated for a service life of 20,000 hours and the fracture/separation occurred after about 8,077 hours. The manufacturer stated that prior to this event there had been no reported similar main rotor blade failures.
HISTORY OF FLIGHT On July 7, 2007, about 1651 eastern daylight time, a Eurocopter EC 130 B4 helicopter, N453AE, registered to Meridian Consulting Company, Inc., and operated by Liberty Helicopters, Inc., sustained substantial damage following an in-flight separation of a section of one of the main rotor blades and subsequent autorotation onto the Hudson River, New York, New York. Visual meteorological conditions prevailed, and a company visual flight rules (VFR) flight plan was filed for the flight, which was operating under the provisions of Title 14 Code of Federal Regulations (CFR) Part 91 and 136 as a revenue sightseeing flight. The flight departed from the West 30th Street Heliport (JRA), New York, New York about 1643. There were no injuries to the certificated commercial pilot or seven passengers. The pilot stated that she had flown the accident helicopter on three previous flights that day and reported no discrepancies on those flights. The accident flight departed for a 10-minute sightseeing flight. All passengers were wearing inflatable life vests, which were contained in a pouch that was strapped around their waists. Approximately 8 minutes into the flight, while flying southbound on a left base leg, or approximately 1/2 mile from the point where she would have turned onto final approach for the heliport, she heard a loud bang, and felt an abnormal vibration (medium to low). At that time, she said she was flying about 350 to 400 feet above the water and between 100 and 110 knots indicated airspeed. The bang was the first thing that got her attention, and noted there was no yawing motion associated with the noise. She saw gray colored debris that was rectangular in shape, and approximately 8 inches in length, fly from the aft left to front left, before it went out of sight. She then heard a "winding down" of the main rotor rpm, but did not hear the low rotor warning horn. The main rotor rpm decay was immediate. She looked at the dual tachometer, but did not recall the main rotor rpm reading. She entered autorotation by applying down collective and aft cyclic, and also deployed the pop-out floats. The vibration was prominent and abnormal. She made a slight flare at 25 feet, and the helicopter settled from that altitude. She applied forward cyclic and the helicopter landed "soft" on the choppy water. The helicopter was level at the point of touchdown and at that time, she had one-half “up” collective applied. She reported there was no binding of the flight controls from the time of hearing the sound to the point of touchdown on the water. After touchdown on the Hudson River, she noted that the main rotor blades were tilted to the right, the tail rotor was still spinning, and the engine was still running. She did not report hearing any horn or seeing any lights, and she did not make any radio calls. She helped the front seat passengers remove their restraints and exit the helicopter. One passenger in the rear seat helped the other rear seat passengers exit the helicopter. All occupants were rescued from the water by private boaters. A witness on a boat reported suddenly hearing a very loud banging noise. The banging noise continued until the main rotor blades contacted the water and became damaged along with what appeared to be pieces of the engine cowling. The banging sound decreased, but the engine remained running “very smoothly,” though it appeared to him to be out of its normally installed position. He said the banging sound as being a thumping sound, metallic in nature, that in his opinion was consistent with the main rotor blades contacting something metallic. The witness, who is an airplane mechanic, reported that the sound was consistent with the sound of tapping of a hard plastic screwdriver handle on aluminum skin. PERSONNEL INFORMATION The pilot, age 37, held a commercial pilot certificate with rotorcraft helicopter and instrument helicopter ratings, issued September 27, 2006. She also holds a private pilot certificate with an airplane single-engine land rating. She was issued a second-class medical certificate with no limitations on February 13, 2007. There was no record of any previous accidents or incidents or enforcement actions by the Federal Aviation Administration (FAA). The pilot was hired by Liberty Helicopters, Inc., on February 8, 2007. Just prior to employment, she reported having a total time of 2,286 hours, of which 1,103 hours were in rotorcraft and 1,183 hours were in airplanes. Three of the 1,103 hours in rotorcraft were in Aerospatiale helicopters. Her last airman competency/proficiency check in accordance with 14 CFR Part 135.293 titled, “Initial and recurrent pilot testing requirements” and also 14 CFR 135.299 titled, “Pilot in command: Line checks: Routes and airports” was performed on February 8, 2007. The flight duration was recorded to be 1.0 hour and the results were listed as “Approved.” The flight was in the accident helicopter. She was qualified to act as pilot-in-command (PIC) in the following make and model helicopters: Eurocopter EC 130 B4, Aerospatiale models AS350B1, AS350B2, and AS350BA. Since being hired by Liberty Helicopters, Inc., the pilot recorded approximately 239 hours in various make and model helicopters including time spent in flight training. She reported on the NTSB Pilot/Operator Aircraft Accident/Incident Report having a total time of 2,752 hours, of which 1,569 were in rotorcraft. In the previous 90 days she reported accruing 357 hours in rotorcraft, of which 214 hours were in the accident make and model helicopter. AIRCRAFT INFORMATION The helicopter was manufactured in December 2001, by Eurocopter as model EC 130 B4, with serial number 3487. It was equipped with an Arriel 2B1 engine, rated for 5 minutes at 747 shaft horsepower. Main rotor blades part number (P/N) 355A-11-0030.00, serial numbers (S/N’s) 22312, 22716, and 22741 were installed at the time of manufacture. Following manufacture, the helicopter was disassembled, shipped to Liberty Helicopters, Inc., and reassembled on January 3, 2002. At that time the main rotor blades were installed in accordance with (IAW) the maintenance manual. A Standard Airworthiness Certificate was issued on April 9, 2002. The type certificate data sheet indicates that for the accident make and model helicopter, the maximum number of passengers is 6. On September 27, 2002, the helicopter was modified IAW Service Bulletin (SB) 25.028 which allowed the installation of 8 seats. The helicopter was placed on Liberty Helicopters, Inc., Operations Specifications on April 22, 2002, and was maintained in accordance with a Federal Aviation Administration (FAA) Approved Aircraft Inspection Program (AAIP). With respect to the airframe, the following inspections are required: 3-day check, 100, 500, 1,000, 2,000, and 2,500-hour inspections. The 3-day check stipulates that the main rotor blades are to be checked for security and general condition of the skin, trim tabs, and polyurethane protective strips. A visual inspection of the main rotor blade for scratches, cracks, impacts and distortions is also indicated. Following the 3-day check, a caution indicates, “Ensure all cowlings and fairings are closed and latched.” Review of the 100-hour inspection checklist revealed that the skin, and leading edge of the main rotor blades, are to be checked for delamination and cracks. The inspection of the blades is accomplished IAW the manufacturer’s aircraft maintenance manual (AMM) 62-11-00, section 6-1. Although the manufacturer Master Servicing Recommendation (MSR) manual specifies to inspect the main rotor blades at intervals of 110 hours, the AAIP work card specifies that the blades are to be inspected for cracks at intervals of 100 hours. Review of the maintenance records revealed the helicopter had a 100-hour inspection on June 23, 2007. The helicopter total time at the time of the inspection was 7,992.5 hours. Additionally, the 3-day check was signed off as being complied with on July 7, 2007. The helicopter had been operated for approximately 85 hours since the last 100-hour inspection, and its total time at the time of the accident was approximately 8,077 hours. On the day of the accident, prior to its first flight, the helicopter and engine total times were recorded to be 8,072.2 and 7,852.7 hours, respectively. The engine was started about 0848, and remained running throughout the day until after the accident. Excluding the accident flight, the helicopter was operated on 25 flights for a total of 4.95 hours. Further review of the maintenance records revealed there was no record that any of the main rotor blades had been repaired or replaced since the helicopter was manufactured. Additionally, there was no record of replacement or major repair to any of the cowlings or fairings. Main rotor blade P/N (P/N) 355A-11-0030.00, has a 20,000 hour service life limit. Eurocopter personnel reported that prior to the accident, there has not been one reported failure of the blade. METEOROLOGICAL INFORMATION A surface observation weather report taken at La Guardia Airport, New York, NY, at 1651, or the approximate time of the accident, indicated broken clouds existed at 7,500 and 25,000 feet mean sea level, the visibility was 10 statute miles, the temperature and dew point were 87 and 53 degrees Fahrenheit respectively, and the altimeter setting was 29.82 inches of Mercury. FLIGHT RECORDERS The helicopter was equipped with dual channels/modules Vehicle and Engine Management Display (VEMD), which is designed to store maintenance data, and is installed in the instrument panel. It displays vehicle and engine parameters, the computation and display of engine limitations, the fail management procedures, the computation and display of weight related to performance data and the number of engine cycles. Relevant VEMD stored data includes flight report, failure report, and over-limits report, which are not dated but time stamped relative to the start of a flight. The helicopter was also equipped with dual channels/modules Digital Engine Control Unit (DECU), which is intended for maintenance use only and records only when specific engine control system discrepancies are encountered. WRECKAGE AND IMPACT INFORMATION The helicopter landed on the Hudson River, and while the fenestron was not attached nor located when the helicopter was recovered, a surveillance video depicted the fenestron attached while the helicopter was upright on the water before recovery. Examination of the helicopter following recovery from the water revealed all main rotor blades remained attached. The fenestron, horizontal stabilizer assembly, long tail rotor drive shaft and landing gear fairings were separated and not recovered. The rear chamber of the left float was found deflated; all other chambers of both floats were found inflated. The engine was rotated approximately 90 degrees to the left and was resting on the left side of the fuselage to attach area. Heat damage to the left side of the tail boom, adjacent to the resting position of the exhaust duct was noted. The main gear box was resting on its right side, displaced aft approximately 10 degrees, and was nearly horizontal. The main gear box crossbeam remained connected to the main gear box at all four attach points and also at the deck attach points; sections of main gear box deck structure remained attached. Three of the four main gear box support struts had evidence of bending overload about the same length, while the right rear main gear box support strut was fractured in the threaded area of the rod end adjacent to the main gear box attach point. The crossbeam with attached main gear box deck structure and the fractured right rear main gear box support strut were retained for further examination by the NTSB Materials Laboratory. The tail boom was fractured approximately 2.5 feet aft of the battery. The right side of the tail boom was compressed/flattened consistent with contact by a main rotor blade. The aft tail rotor drive shaft cover was separated and not recovered, and the tail rotor flex ball cable was fractured in the area of the tail boom fracture area. Continuity from the anti-torque pedals to the fracture point of the flex ball cable was confirmed. Cyclic and collective control continuity was confirmed from the cockpit to the main gear box deck area; bending overload was noted at the flight control separation points. There was no evidence of any preexisting failure or malfunction of the cowling, cowling latches, main gearbox, or tail rotor drive system. Examination of the main rotor blades with NTSB oversight was done by representatives of American Eurocopter, Eurocopter, the FAA, and Turbomeca USA. The examination of blade S/N 22312 revealed a separation of material aft of the spar. All three main rotor blades were sent to the manufacturer's facility in France for further examination. Inspection of the cockpit and cabin revealed no deformation of the cockpit or cabin floor. A total of seven inflatable life vests were found. No damage was noted to any of the eight seats; all seats restraints operationally checked good during postaccident pull testing by hand. The VEMD and DECU were retained for further examination. Examination of the engine at the manufacturer’s U.S. facility with NTSB oversight revealed no evidence of preimpact failure or malfunction of the engine or engine accessories. Examination of the power turbine wheel assembly revealed 20 of the 37 blades appeared ruptured at the design shear point (consistent with separation due to overspeed). The fractured surfaces appeared granular in nature. The shaft and both sides of the power turbine wheel assembly exhibited significant rotational scoring. Approximately two-thirds of the forward edge of the power turbine disk appeared ground down, and the forward portion of the teeth of the power turbine nut appeared flattened and smeared. The power turbine shroud (containment ring) appeared elongated and exhibited marring consistent with power turbine blade separation. MEDICAL AND PATHOLOGICAL INFORMATION The pilot submitted a urine specimen on the date of the accident for post accident drug testing. The results were negative for tested drugs consisting of amphetamines, cocaine metabolites, marijuana (THC), phencyclidine (PCP), and opiates. TESTS AND RESEARCH The NTSB Materials Laboratory Factual Report indicates that with respect to the examination of the structure attached to the crossbeam, the fracture surfaces showed deformation and fracture features consistent with overstress fracture. The examination of the fractured right rear main gear box support strut revealed the fracture was consistent with overstress fracture. The examination of the laminated disks revealed that some exhibited bulging, but no other visible signs of degradation or defects were noted. Readout of the VEMD performed by the Bureau d’Enquêtes et d’Analyses (BEA) revealed that flight Nos. 7178 and 7179 were recorded by both channels (modules) during the accident flight. The flight duration value (Ng value greater than 50 percent) was 8 hours 23 minutes; while the engine had been operated for approximately 8 hours 3 minutes at the time of the accident. A total of 9 failures were recorded during flight 7178, and 6 failures were recorded during flight 7179. Two over-limits were recorded during flight 7178, and no over-torque was recorded during flight 7178 or 7179. The “free turbine” value translated to NR speed was recorded to be 504 rpm, while the normal maximum rpm at sea level is 399 rpm. A total of 4 minutes elapsed between the Ng value dropping below 50 percent and VEMD shutdown. The recorded failures were “CROSS FQ”, “TEST FADEC 4”, “SURV DOM EOT”, “SURV UNDOFF NR”, and “TEST FADEC 7”. A minimum of 17 minutes 26 seconds passed between the first recorded failure and the shutdown of the equipment. During that period, several failures occurred related to the collective pitch and leading amber GOV lighting on the annunciator panel. Readout of the DECU by the BEA revealed a total of six failures were recorded during the accident flight. One failure related to fli
The fatigue fracture and in-flight separation of a 20-inch section of the blue composite main rotor blade trailing edge, aft of the spar, due to inadequate manufacture, and the manufacturer’s subsequent failure to detect an out-of-specification deviation in the rotor blade’s trailing edge roving.
Source: NTSB Aviation Accident Database
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