NTSB Identification: MIA07FA116

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.


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.


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.


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.


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.


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.


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.


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 flight that was recorded by both channels/modules indicated, “PANNE_XPC_OUI and CAPTEUR_XPC.” The message indicates a discrepancy has been encountered in the collective pitch signal. The signal from the collective pitch potentiometer was found to be out of limits (in terms of maximum or minimum reached value or in terms of signal variation). This leads to an amber “GOV” illuminating on the helicopter failure panel and sets a backup collective pitch value. In this degraded mode the engine remains in automatic mode, and the flight manual cautions the pilot against sudden variations of the collective pitch.

The main rotor blade, P/N 355A-11-0030.00, is composed of a spar, skin, foam filler, trailing edge roving, and trailing edge tab. The spar is made from a glass fiber reinforced roving wound about polyurethane foam. The skin and trailing edge tab are made of glass fiber reinforced fabric, and the trailing edge roving also has glass fiber reinforcement. The foam filler is located between the upper and lower skin and between the spar and the trailing edge roving. From the root-end area to blade station 3200, the skin is reinforced by additional layers of glass fabric. The width of the trailing edge tab is less at blade station 1350 relative to blade station 1100. Correspondingly, the position of the trailing edge roving located at the leading edge side of the trailing edge tab shifts closer to the trailing edge between blade stations 1100 and 1350. Additionally, the trailing edge of the foam core is shifted closer to the trailing edge at blade station 1350 relative to blade station 1100, to fill space left by the trailing edge roving moving towards the trailing edge.

Eurocopter reported that main rotor blade S/N 22312 was manufactured in February 2001, and the maintenance records reflect that the blade was installed on the helicopter in November 2001. It remained installed on the helicopter until removal for shipping of the helicopter to the United States. The main rotor blade was reinstalled on January 3, 2002. There was no record that the blade had been repaired; the total time on the blade was approximately 8,077 hours.

Examination of the main rotor blade S/N 22312 at the manufacturer’s facility with BEA oversight revealed that a section of blade aft of the spar from blade station 1308 to the blade tip was separated. An undulation in the glass fiber reinforced trailing edge roving between blade stations 1288 and 1308 (0.79 inch long) was noted during x-ray inspection of the blade. The undulation was at the end of a skin reinforcement layer. The undulation, which occurred during blade manufacture was not in accordance with the manufacturing specification (the glass fibers of the trailing edge roving should have been nearly straight and parallel with the blade span). Further examination of blade station 1300 revealed part of the trailing edge roving was displaced toward the leading edge into the space of the foam filling and toward the trailing edge into the space of the trailing edge tab. A cavity and a resin-rich area were noted in the trailing edge roving.

The blade was fractured chordwise at blade station 1308; the fracture surface of the trailing edge roving exhibited flat fiber fracture and pitting consistent with compression. The fracture features of the glass fiber reinforced skin at blade station 1308 exhibited flat fracture at 90 degrees relative to the direction of the fibers, which is consistent with fatigue. The fatigue area was localized to within 1.18 inches (30 millimeters) from the trailing edge. Beach marks consistent with fatigue crack propagation were noted in the resin between the glass roving and the trailing edge tab on several small areas of the fracture surface, each measuring approximately 0.001 square inch in area.

The edges of the fracture surface at the trailing edge exhibited a rounded pattern instead of being square to the surface and exhibited evidence of erosion consistent with a preexisting crack. The direction of skin crack propagation beginning at blade station 1308 was from the trailing edge to the spar, then outboard towards the tip to approximately blade station 1800, then back to the trailing edge. The remainder of the skin separation from blade station 1800 to the tip occurred as the result of water contact. The orientation of the skin and reinforcement fabrics and the percentage of resin in the skin and reinforcement fabrics on the upper and lower surfaces were correct.

Pieces from the trailing edge of the yellow and blue main rotor blades were submitted to the National Transportation Safety Board's Materials Laboratory for further examination. Examination of the trailing edge roving associated with the blue main rotor blade revealed fiber ends with chop marks consistent with local fiber bending fractures under compression or transverse shear loading. The compression sides of the fiber fractures were generally located at the lower half of the fibers. On the tension sides of the fibers, hackle regions were observed with features emanating radially from the origin at the upper sides of the fibers.

Areas of the trailing edge roving had fiber ends that appeared relatively smooth with no hackle regions. Many fibers were smooth and flat across the entire diameter, while other fibers had a change in fracture plane at one side or opposite sides of the diameter appearing similar to a chop mark offset toward the surface of the fiber. An area where fatigue features were identified in the resin near the trailing edge end of the trailing edge roving was noted. A bundle of skin fibers were located adjacent to the resin fatigue region; the fibers in this bundle of fibers generally had relatively smooth fracture features with little or no hackle region.

The trailing edge roving portion was examined using scanning electron microscopy (SEM). The fracture surface had many broken fiber fragments and generally appeared damaged from post-fracture contact. Many fiber ends were covered with matrix material and/or deposits that obscured the fiber fracture surfaces. Where fiber ends were visible, the fractures generally showed a mix of features with some areas having fibers with chop marks near the middle of the fiber and hackles on the tension side and other areas having fibers with relatively smooth fracture features with no hackles. Chop marks are step changes in the fracture plane on the fiber fractures near the middle of the fiber that occur as a result of local bending of the fiber. Hackle regions are step features that radiate outward parallel to the direction of fracture propagation resulting from separating portions of the crack surface that are parallel but not coplanar.

An SEM view of an area of lower skin fibers on piece identified as "C" near the leading edge end of the trailing edge roving revealed the fibers were all fractured in nearly the same plane and showed smooth fracture features with no hackle regions or chop marks, features consistent with fatigue cracking. The fracture surface had a roughened appearance with curving lines perpendicular to the fracture direction, features consistent with fatigue. A close view of fractured fiber ends in the skin in piece identified as "E" revealed fracture features on the fiber ends were relatively rough and generally showed hackle features radiating outward consistent with overstress fracture of glass fibers under tension loading.

Radiating outward from the origin, glass fractures can have mirror, mist, and hackle regions. The mirror region is a relatively smooth region that surrounds the origin. The mist region has a misty appearance under optical magnification and is an indication of the crack front beginning to break up as the crack approaches terminal velocity during fracture. The hackle region is located outside of the mist region and occurs when the crack front becomes discontinuous and propagating on different planes during fracture. Hackle regions can be useful for determining the size of the mirror region. Mirror and/or hackle regions were identified on the glass fiber fracture surfaces. In many fibers of the trailing edge roving, the mirror region extended across the entire fiber diameter.

Empirical studies have shown the size of the mirror region can be used to calculate the stress on the glass at the time of breakage. The mirror radius, R, is a measure of the size of the mirror region, and the stress, s, on the fiber at the time of fracture is inversely proportional to the square root of the mirror radius. For example, a larger mirror radius corresponds to a lower stress at the time of fracture.

Damage on the trailing edge of the yellow blade was examined optically and using SEM. Optically dark areas on the damage surface showed a high peak of aluminum when analyzed using energy dispersive x-ray spectroscopy consistent with metal transfer from contact with an aluminum object. Smearing and damage patterns were consistent with sliding contact with an object moving relative to the blade in a
direction toward the trailing edge of the blade.

Eurocopter personnel estimated that the separated section of skin between blade stations 1308 and 1830 accounted for a loss of approximately 400 grams. They also performed calculations which indicated that in the event of an imbalance generated by a loss of 400 grams from one main rotor blade at 1555 mm from the blade root, the strength of the main gear box support struts is assured for all modes of deterioration including static, buckling, and fatigue. Eurocopter also performed analysis pertaining to separation of a section of main rotor blade and although the document was considered proprietary, the conclusion was that there would be no impact on control positions, aircraft attitudes, speeds, and general aircraft response to control inputs.

Inspection of main rotor blade S/N’s 22716, and 22741 at the manufacturer’s facility with BEA oversight revealed no evidence of out of tolerance discrepancies pertaining to blade station 1300. The trailing edge of blade S/N 22741 at blade station 1167 (approximately 6 inches from blade station 1308) exhibited an impression. Traces of aluminum were noted on the surface of the impression but no determination could be made as to what caused the impression.

Review of the documents associated with manufacturing of main rotor blade S/N 22312 revealed that during a quality control visual inspection following molding, an additional x-ray inspection was requested of blade station 1300 in an effort to determine whether any materials were “locked up” between the two outboard skins. The x-ray inspection of blade station 1300 was performed, and the review of the x-ray by the technician revealed the initial remark to be minor in nature and the production process of the blade was completed. No further discrepancies associated with manufacturing of the blade were reported.

During the investigative process, Eurocopter personnel inspected their manufacturing files for a similar entry concerning a quality control request for an additional x-ray inspection of blade station 1300. Their review of 7,991 files revealed only one record pertaining to main rotor blade P/N 355A11.0030.04, S/N 28427, which was manufactured on February 3, 2006. The manufacturing files indicated a quality control request for an additional x-ray of the trailing edge area around station 1300. The x-ray inspection occurred and no defects were found. The blade was placed into service and subsequently located in Papouasy, New Guinea. The blade, which had not accumulated any time since its manufacture, was also sent to Eurocopter for examination with BEA oversight. The examination revealed no out of tolerance discrepancies relative to manufacturing specifications, specifically at blade station 1300.

Following the accident, the helicopter manufacturer prepared a TELEX alerting operators of specified model helicopters to inspect the trailing edge skin of main rotor blades.