WPR10FA133
WPR10FA133

HISTORY OF FLIGHT

On February 14, 2010, about 1505 mountain standard time, a Eurocopter Deutschland GmbH EC135 T1 twin-engine turbine-powered helicopter, N127TS, lost control and impacted terrain near Cave Creek, Arizona. The pilot and four passengers were fatally injured. The helicopter was substantially damaged. The helicopter was registered to Services Group of America Inc. (SGA), of Scottsdale, Arizona, and was operated under the provisions of 14 Code of Federal Regulations Part 91 as a personal cross-country flight. Visual meteorological conditions prevailed at the time of the accident, and no flight plan was filed. The flight departed the Whispering Pines Ranch, Parks, Arizona, about 1430 and was destined for Scottsdale Airport (SDL), Scottsdale, Arizona.

SGA personnel reported during postaccident interviews that the helicopter made frequent trips between SDL and the Whispering Pines Ranch and that the helicopter had arrived at the ranch on February 12, 2010. In a statement submitted to the National Transportation Safety Board (NTSB) investigator-in-charge (IIC), the ranch foreman indicated that he had spoken to the helicopter owner while the helicopter was en route to the ranch on February 12. He was not present when the helicopter arrived and did not know who was on the helicopter when it arrived other than the pilot and owner, but he recalled helping the pilot put the helicopter in the hangar after its arrival.

In a postaccident interview with the IIC, the ranch foreman reported that the helicopter was scheduled to return to SDL on the day of the accident. The foreman stated that, on the morning of the accident, the pilot, four passengers, and two dogs arrived at the helicopter hangar for the flight to SDL. The foreman revealed that he placed the passenger’s personal items on the ground outside of the baggage compartment, which was located at the rear of the helicopter; he said that the pilot always loaded the baggage compartment himself. The foreman reported that after the pilot had completed loading the helicopter and doing his preflight inspection, he entered the helicopter, sat in the right front cockpit seat, and started both engines. The foreman added that he then assisted the adult female into the rear cabin area, where she occupied the left rear aft-facing seat. Subsequently, a male passenger boarded the helicopter and sat in the right rear forward facing seat. This was followed by a small dog that occupied the left rear forward-facing seat directly across from the female passenger and a larger dog that was positioned on the floor between the left rear aft-facing and forward-facing seats.

The foreman reported that after the male and female passengers and the two dogs were on board, he closed the right passenger door and ensured that it was locked. The foremen further reported that he moved forward to a position that was immediately outside of the right front pilot's position. He observed the helicopter owner and his 5-year-old, 42-pound daughter walk around in front of the helicopter and board the helicopter from the left forward cockpit door where they both occupied the left front cockpit seat, with the small girl positioned on her father’s lap. When asked how frequently the child occupied the left front cockpit seat with her father, the ranch foreman replied "occasionally." The foreman stated that he could not tell if either the helicopter owner or the child were secured and restrained in the helicopter. The foreman revealed that on previous flights, the helicopter owner had strapped his daughter in on top of him. He said that after everyone was on board, he closed the right passenger door and ensured that it was locked and secured. He said he then went forward to the right front cockpit window area, looked at the pilot, who had his shoulder harness and seatbelt on, and motioned to him that the right passenger door was secured and the helicopter was ready for departure. The foreman indicated that he then proceeded away from the helicopter to his pickup truck, which was positioned about 90 degrees to the right (west) of the helicopter, which was oriented to the south. The foreman said that from his truck, he watched the helicopter lift off, ascend to about 100 to 150 feet, then make a 180 degree turn to the north and begin forward flight, after which the helicopter departed to the northwest.

Several witnesses to the accident were either interviewed by or submitted written statements to the IIC.

Witness #1, who was located on the helicopter's flightpath and was about 1,385 feet north of the accident site, reported that while facing east, he heard two pops. The witness stated that he subsequently looked to the southeast where he estimated the helicopter was about 300 feet above ground level (agl); at the same time, he observed blade debris separate from the rear of the helicopter. The witness added that the helicopter then turned to the west before going inverted and appeared to go straight down. The witness added that the main rotor blade was making a loud clapping sound.

Witness #2, who was located about 500 feet west of the helicopter's flightpath and was about 2,000 feet north-northwest of the accident site, reported that he heard the helicopter approaching from the north and that the engine was cutting out; he described it as popping a couple of times. The witness stated that he observed the helicopter in a rotation but could not indicate in which direction. He further stated that it went around in a circle two times and then suddenly went down in a steep angle of about 30 degrees while cork screwing. The witness added that he did not observe any debris separate from the helicopter.

Witness #3, who was in line with the helicopter's flightpath and was about 2,350 feet due north of the accident site, stated that he initially saw the accident helicopter when it was about 1 mile north of his position and that it sounded perfect when he first sighted it. The witness reported that, about 10 seconds later, he heard what sounded like two small pops, followed by the helicopter making roaring noises. He said he then observed two flashes on the top of the main rotor. The witness stated that the helicopter started spinning and losing altitude and that it spun at least three times; he indicated that he then heard the engine cut out and make a big pop. The witness added that the tail of the helicopter went down, the nose went up, and then the aircraft fell tail first. In a follow-up interview, the witness reported seeing the helicopter spiral in a circular motion towards the ground and then rapidly gain altitude before it flipped upside down and spiraled nose first into the ground.

Witness #4, who was inside of his home and was about 1,250 feet east southeast of the accident site, reported that he heard a loud and unusual sound. The witness stated that he then went outside and observed the helicopter go straight down from about 200 feet before it went out of view.

Witness #5, who was about 500 feet west of the helicopter's flightpath and was initially in his house about 1,000 feet north-northwest of the accident site, stated that he first heard a rotor noise but that the noise changed, and then he heard a pop, followed by popping and banging sounds for a few seconds. The witness reported that he then went outside and observed the helicopter go down. He described the descent as a nose dive, estimated to be at about an 80 degree nose-down attitude before it went out of sight.

Witness #6, who was about 650 feet east of the helicopter's flightpath and was about 2,000 feet north-northeast of the accident site, reported that he heard a popping sound. The witness stated that he observed the helicopter when it was about 400 feet agl moving to the south and that he heard it making popping sounds about every 15 to 20 seconds. The witness stated that he observed the helicopter make four circles in a level, clockwise direction before it went down and out of site.

Witness #7, who was at the same location as the sixth witness, stated that he saw the helicopter overhead as he was looking to the west. The witness reported that he heard a popping sound every 15 to 20 seconds and then observed the helicopter rotate clockwise two or three times, then nose down slightly. The witness added that the helicopter then rolled to the left, went nose down, rolled to the right nose down again, and impacted the terrain.

Witness #8, who was about 1,000 feet west of the helicopter's flightpath and located about 1,500 feet northwest of the accident site, reported that he noticed the helicopter as it approached from the north and that, after a few seconds, he heard a loud pop and observed the helicopter spinning and losing altitude. The witness added that he observed the helicopter make several rotations before going out of sight; the witness was not certain of the direction of the rotations.

Witness #9, who was about 3,000 feet west of the helicopter's flightpath and was about 3,300 feet west-northwest of the accident site, reported that he observed the helicopter when it was traveling from north to south; he estimated it was about 1/2-mile east of his position. The witness reported hearing what he described as a noise like the engine was rapping and that the helicopter was maybe 300 to 400 feet high. The witness added that when the helicopter was about one-half mile east of his residence, there was a system failure of some kind, which made a loud noise with parts observed separating from the helicopter. The witness stated that at this time, the aircraft made 3 or 4 clockwise rotations, followed by the nose of the helicopter nosing over between 25 to 45 degrees, after which the helicopter disappeared from sight.

Witness #10, who was about 1.34 miles west-southwest of the main wreckage site, reported in a statement submitted to the IIC that she initially became aware of the helicopter because of the noise it was making, such as popping sounds and blades that sounded louder than normal. The witness stated that the blades were making loud "whop, whop" type sounds they usually do when they are taking off or landing and that some of the noise sounded like a backfire. The witness reported that the helicopter was hovering perfectly still for a short period of time, about 10 seconds, and that its nose was pointing basically to the north, which gave her a clear view. The witness further reported that the helicopter was above the horizon of the mountains and that it was not losing altitude. The witness added that as she continued to watch the helicopter, it suddenly dove straight down to the ground, about half way, before it pulled up and traveled to the south. The witness revealed that at this time, it started to go in tight circles and circled two and a half times before it dropped suddenly and disappeared from view.

The helicopter wreckage was located just north of a river wash about 10 nautical miles (nm) north and in line with its intended destination (SDL). First responders reported that no postcrash fire was detected upon reaching the accident location just a few minutes after the helicopter crashed. However, about 2 minutes later, a small fire erupted, which resulted in a larger, intense fire that consumed the helicopter.

PERSONNEL INFORMATION

Pilot-in-Command (right forward cockpit seat occupant)

General

The pilot, age 63, possessed a Federal Aviation Administration (FAA) commercial pilot certificate for rotorcraft-helicopter, which was issued on December 6, 1969. The pilot’s most recent second-class FAA medical certificate was issued on July 31, 2009, with the limitation that he "must wear corrective lenses and possess glasses for near and intermediate vision." The pilot reported a total flight time of 10,117 hours on his most recent airman medical application. SGA flight operations personnel reported that the pilot had accumulated 11,045 total flight hours, 824 hours in the EC135 T1, and that he had flown 13 hours in the preceding 90 days. It was also reported by SGA that the pilot's most recent flight review was conducted on July 1, 2008, in the EC135 T1.

SGA flight operations personnel revealed that the pilot was a U.S. Army helicopter pilot in Vietnam; however, no record of military flight time was obtained during the investigation. Further, the pilot’s personal pilot logbook was not obtained during the course of the investigation.
EC135 T1 flight training documentation

A review of the pilot's Eurocopter EC135 T1 training records revealed that the pilot received training at the American Eurocopter training facility in Grand Prairie, Texas, and additional training offsite at the customer’s facility. The pilot’s initial training consisted of transition ground school training in the EC135 T1, which was completed on June 28, 2002, and included 7.2 hours of flight training. The pilot subsequently completed recurrent ground school training on September 11, 2003, with 4.3 hours of flight training; on April 26, 2004, with 1.5 hours of flight training; on May 17, 2006, with 2.6 hours of flight training; and on July 1, 2008, with 3.2 hours of flight training. All ground and flight training records indicate satisfactory performance with no deficiencies noted.

American Eurocopter flight instructor statements

During the initial phase of the investigation, two American Eurocopter instructor pilots, both of whom had provided the accident pilot with EC135 T1 instruction from 2002 through 2008, submitted statements regarding concerns they had with statements the accident pilot had made during training.

The first instructor pilot provided the accident pilot with his initial transition ground school training in 2002, followed by recurrent training in 2003, 2004, and 2006. The instructor pilot reported that during his training in 2002, which was conducted in Seattle, Washington, the accident pilot displayed an abnormally high degree of pressure to accomplish flights from the helicopter’s owner and that he was visibly shaken when discussing the amount of pressure he received. The instructor stated that during the week he spent training the accident pilot, the conversation regarding his employer often turned to the difficulties he endured to keep flights on schedule. The instructor pilot further stated that one conversation he had with the accident pilot characterized the amount of pressure that was present to complete missions. The instructor revealed that the accident pilot stated that it would not be uncommon to fly the helicopter’s owner from Seattle to his home on Vashon Island when the weather conditions at night were so poor that they would follow the ferryboat lights to navigate across the bay under foggy conditions.

The first instructor pilot also reported that in 2004, he provided the accident pilot recurrent training after an incident that damaged the helicopter. The instructor further reported that the accident pilot stated that he was to fly to Vashon Island, Washington, to pick up the owner’s wife to fly her to the Seattle airport. After landing [at Vashon Island], the accident pilot left the helicopter’s engine running and the controls locked while he loaded the passengers and bags. When he attempted takeoff, the cyclic control lock was still engaged, which resulted in damage to the tail boom following the attempted landing. The instructor pilot added that the accident pilot admitted to him that he was flustered because he had to hurry and depart as soon as possible.

The second instructor pilot, who provided instruction to the accident pilot during the summer of 2008, stated that he remembered the pilot commenting about how the helicopter's owner dominated the cockpit duties before a flight. The instructor added that the accident pilot revealed that when the owner (who was not a rated helicopter pilot but was a rated fixed-wing airplane pilot) flew, he would get in the cockpit, flip switches, and go. The instructor reported that he felt that the accident pilot was intimidated by the owner and would not insist that proper aircraft procedures be followed.

In a submission to the NTSB IIC, the SGA chief pilot for flight operations characterized the accident pilot as a Vietnam-era combat pilot who had also flown for the U.S. Forest Service in the Pacific Northwest Region, a very demanding region of the United States in which to fly. Additionally, he characterized the accident pilot as a conscientious and professional pilot in every sense of the word who would not be intimidated.

Helicopter Owner/Certified Airplane Pilot (left forward cockpit seat occupant)

A review of FAA records revealed that the helicopter's left forward cockpit seat was occupied by the helicopter’s owner, age 64, who possessed a private pilot certificate for airplane single engine land, which was issued on October 26, 1967. A further review of FAA records failed to reveal any documentation to indicate the issuance of an airman medical certificate. Additionally, the investigation failed to reveal any background information about the pilot's aviation background, such as pilot training or total flight hours, since no pilot records were recovered during the investigation. It was also revealed that the pilot was not rated to fly helicopters. In a February 19, 2010, interview, the SGA chief pilot, when asked if the helicopter's owner ever flew the helicopter, responded that it was common for him to fly. The chief pilot further commented, "He liked to fly."

AIRCRAFT INFORMATION

The accident helicopter, serial number (S/N) 0094, was manufactured in 1999 by Eurocopter Deutschland GmbH and received its FAA standard airworthiness certificate on April 21, 2000. According the maintenance records, on the day of the accident and before departing from the Whispering Pines Ranch, the helicopter had accumulated a total time since new (TTSN) of 1,115.9 hours. The accident flight time was estimated to be about 0.7 hours, which would have resulted in a total flight time at the accident site of 1,116.6 hours TTSN.

A review of maintenance records for N127TS revealed that the helicopter was maintained in accordance with Eurocopter's recommended inspection program. The most recent annual inspection was conducted on October 30, 2009, at a total airframe time of 1,103 hours. No discrepancies were noted during the records review.

SGA purchased the helicopter from the original private owner in 2002. The helicopter had previously been operated from a luxury yacht and was configured with a utility interior; SGA subsequently configured the helicopter with an executive interior. This configuration included the pilot and copilot seats in the front, two aft-facing seats and a galley in the center of the cabin, and two forward-facing aft seats.

Previous incidents involving N127TS

On May 8, 2003, while on final approach to Vashon Island, Washington, the helicopter owner, who was not a rated helicopter pilot, was at the controls in the left front cockpit seat. Reportedly, the left seat was not in the proper detent position and slid aft. The helicopter dropped about 50 feet but was recovered by a quick collective input. In an incident report submitted by American Eurocopter, it was reported that a loud bang was heard, followed by the touchdown of the helicopter. A postaccident examination of the helicopter revealed impact damage to the horizontal stabilizer. Additionally, numerous pieces of the engine were located on the ground behind the helicopter. The helicopter was repaired at both the facilities of American Eurocopter, Grand Prairie, and at the facilities of Cascade Airframe Repair, of Seattle, Washington. The helicopter was returned to service on August 13, 2003.

On January 14, 2004, the accident pilot made a hard landing in the accident helicopter while attempting to land at a grassy heliport located on Vashon Island, Washington. The helicopter was subsequently repaired by American Eurocopter and returned to service on April 24, 2004. The helicopter was piloted by the accident pilot at the time of this event.

On September 13, 2007, shortly after departing Flagstaff, Arizona, on a cross-country flight to Scottsdale, Arizona, while at a cruise altitude of 1,500 feet agl with the autopilot engaged, the #2 engine chip light illuminated, followed by a yaw and an engine shutdown. A single-engine landing was subsequently performed without further incident. The engine was replaced on January 11, 2008, by Cascade Airframe, Seattle, Washington. It was also reported that the helicopter had experienced two #2 engine chip light illuminations in the preceding eight days.

Main rotor blade repair (yellow blade S/N 765)

According to the SGA director of maintenance, main rotor blade S/N 765 (yellow blade) was removed from the accident helicopter in November 2009 and sent to the American Eurocopter facility in Grand Prairie, Texas. A rental main rotor blade (SN 2472, a replacement yellow blade, which was involved in the February 14, 2010, accident) replaced main rotor blade S/N 765. The removal of the S/N 765 blade occurred when maintenance personnel at Cascade Airframe Repair, Seattle, Washington, could not get the blade to balance correctly. The examination by American Eurocopter maintenance personnel revealed a trailing edge crack on both the top and bottom surfaces. The crack was located inboard and found only after the blade had been stripped. It was subsequently determined that the work to repair the blade was beyond the capability of the maintenance facility in Texas, so the blade was sent to the Eurocopter maintenance facility, Eurocopter Deutschland GmbH, on November 13, 2009, for evaluation and repair. The following work was completed on February 1, 2010, at the Eurocopter Germany facility:

• Replaced PU (Polyurethane)-Erosion strips (inside and outside), replaced trim-tabs, replaced auxiliary tab.

• Replaced spherical bearing of support pre assembly, repaired T/E (trailing edge) delamination (trim-tab area).

• Repaired skin delamination of T/E (1 meter rep.) chapter 05-01-04, replaced discharger.

• Repaired skin-core delamination (blade tip area), static and dynamic balanced.

Helicopter General

The EC135 is a light multipurpose twin-engine helicopter. There are five seats in the basic version, which can be extended to eight seats. The accident helicopter had been converted to the 6-seat Executive version.

The main transmission is a two-stage flat gearbox, which is mounted by an antiresonance rotor isolation system on the transmission deck.

The helicopter is equipped with a four-bladed hingeless and bearingless main rotor. The inboard flexbeam enables movement of the blades in all axes. Blade pitch angles are controlled through integrated glass/carbon fiber control cuffs.

The main rotor control linkage system is of conventional design. The hydraulic system for the main rotor controls is designed as a duplex system with tandem pistons (both systems are active). If one system fails, the remaining system has sufficient power to ensure safe flight conditions and a safe landing.

The helicopter is equipped with a Fenestron tail rotor system. There are 10 blades rotating in a housing integrated in the tail boom. The Fenestron is controlled via a Flexball type cable, routed from the pedals to the input control rod of the Fenestron. The tail boom can be separated from the fuselage and consists of a tail boom cone, the horizontal tail plane stabilizer with endplates, vertical fin with integrated tail rotor, and tail rotor gearbox and fairing.

Engine Data

The helicopter was equipped with two Turbomeca USA Arrius 2B1 turboshaft engines, each rated at 577-shaft horsepower.

The #1 Engine (left engine)

The #1 engine (S/N 30167) was manufactured on March 19, 1999. On April 5, 2004, the engine was inspected at the facilities of Turbomeca USA, Grand Prairie, Texas, as a result of a hard landing. At the time of the inspection, the engine had accumulated 452.2 hours TTSN. The most recent inspection of the engine before the accident flight was performed on October 30, 2009, at a TTSN of 1,102.5 hours.

The #2 Engine (right engine)

The #2 engine (S/N 30291) was manufactured on March 20, 2004. On September 19, 2007, at a TTSN of 392.2 hours, the engine was removed and repaired at the Turbomeca USA Grand Prairie facilities for a chip light anomaly. On January 11, 2008, the engine was reinstalled on the accident helicopter (S/N 0094). The most recent inspection of the engine before the accident flight was performed on October 30, 2009, at a TTSN of 227.3 hours.

METEOROLOGICAL INFORMATION

About 1454, the weather reporting facility at the Deer Valley Airport, Phoenix, Arizona, located about 12 nm southwest of the accident site, reported wind calm, visibility 10 miles, sky clear, temperature 23 degrees Celsius (C), dew point -03 degrees C, and an altimeter setting of 29.98 inches of mercury.

COMMUNICATIONS

A survey of air traffic facilities revealed that there were no communications between the pilot and air traffic control (ATC) on the day of the accident.

RADAR DATA

Radar data was received from the Phoenix terminal radar approach control. The helicopter was first observed about 11 miles south of the Whispering Pines Ranch at 1443:30. The helicopter was then observed to be flying south toward SDL at a mode C reported altitude of 7,500 feet mean sea level (msl) until 1451:16, at which time it initiated a gradual descent that lasted until 1457.41; the helicopter was then at 6,000 feet msl. From this point, the helicopter initially climbed and then began a gradual descent with slight altitude, heading, and airspeed fluctuations until 1503.27; the helicopter was then at 4,400 feet msl. Radar data then depicted the helicopter make a rapid climb from 4,400 feet to 4,700 feet within about 5 seconds, followed by a rapid descent to 3,800 feet msl within the next 5 seconds. The last radar hit, which was at 1503:37, revealed that the helicopter’s position was 33 degrees, 50 minutes, 46.67 seconds north latitude and 111 degrees, 55 minutes, 27.81 seconds west latitude, on a magnetic heading of 171 degrees.

Radar Study

The NTSB chief scientist, assisted by an NTSB senior ATC specialist, reported in a radar study that FAA-provided recorded radar data depicted the track of the helicopter from near the departure point to the crash site. However, witnesses to the accident believed that the helicopter passed overhead at about 300 to 500 feet agl, which conflicted with recorded radar data that showed the helicopter to be at 4,400 feet msl, or about 2,000 feet above agl.

Radar data revealed that the helicopter was first observed at an altitude of 7,500 feet msl about 11.5 nm south of the departure point and proceeding in a southerly direction. As the helicopter crossed a ridge (elevation 3,200 feet msl just north of the witnesses), the radar data revealed its altitude was 4,900 feet msl, or 1,700 feet agl. Radar data further revealed that the helicopter then descended to 4,400 feet msl, or about 2,400 agl, followed about four seconds later indicating that the helicopter was at 4,700 feet msl, and 1 second later at 4,400 feet msl. About 4 seconds later, two nearly simultaneous radar returns showed the helicopter was at 3,800 feet msl and 3,700 feet msl. The study indicated that the multiple, mixed mode returns (two radar only returns [skin paint] and one beacon-only return) are consistent with the helicopter remaining at a nearly constant altitude, rather than climbing 300 feet in 4 seconds and descending 300 feet in 1 second. The study further revealed that the 300 feet excursion to 4,700 feet was likely due to the corruption of the static pressure as a result of a severely yawing helicopter. Additionally, the study revealed that radar-only and beacon-only returns often occur when an aircraft is in extreme maneuvers which may blank the transponder. The study added that the last known position was observed over the accident site at 3,700 feet msl.

The main debris field was located directly under the last radar returns from the helicopter. An examination of radar data disclosed no data that would be consistent with the helicopter descending or returning for another pass at a lower altitude. (For more information, see the NTSB's Chief Scientist’s Radar Study, which is located in the public docket for this accident.)

WRECKAGE AND IMPACT INFORMATION

On February 15, 2010, investigators from the NTSB and the FAA, accompanied by representatives from American Eurocopter and Turbomeca USA, examined the wreckage at the accident site. Representatives from the German Federal Bureau of Aircraft Accident Investigation (BFU) and Eurocopter Deutschland joined the investigation on February 18, 2010.

The wreckage was located just north of a river wash on a residential gravel access road about 14 nautical miles north of Scottsdale Airport. The linear debris path extended north from the main wreckage about 0.34 nm, and consisted primarily of pieces of yellow main rotor blade and the left horizontal end-plate. A global positioning system-measured elevation at the accident site of 2,356 feet msl. The main debris field was about 1,795 feet long and about 350 feet wide, on a north to south orientation and in line with the helicopter’s observed flightpath.

The onsite examination of the wreckage revealed that all major aircraft components were accounted for near the wreckage or along the flightpath. There was no evidence of an in-flight fire.

The first identified piece of debris along the flightpath leading to the wreckage site was evidenced by a piece of ribbon from the helicopter’s left vertical endplate, which was located about 1,795 feet north of the main wreckage site. An area of main rotor blade debris that included foam, honeycomb, and paint chips was then identified, which extended from the initial point of the debris field south for about 1,050 feet to a two-lane paved road that was oriented in an east-to-west direction. A large piece of the yellow main rotor blade was located just north of the paved road, which was about 750 feet north of the main wreckage site. A fragment of the helicopter’s tail rotor drive shaft flex coupling was found lying on the south side of the paved road about 695 feet north of the wreckage. A lower piece of the left vertical fin was located about 450 feet north of the main wreckage. A further survey of the area revealed that a piece of the yellow blade tip was found about 400 feet northwest of the main wreckage, with myriad small fragmented pieces of the helicopter observed between the blade tip and the wreckage site. The aft section of the steel tail rotor drive shaft was located about 120 feet northeast of the main wreckage. The main rotor blades and hub found were located about 25 feet north of the primary wreckage site.

The cockpit and cabin areas were destroyed by impact forces and the postcrash fire. The cockpit floor structure was observed to have exhibited accordion compression in the aft direction. The twist grip throttles were both found past the neutral detent and in the high range. The anti-torque pedals were found in the near-neutral position. All flight controls were accounted for, and all flight control tube fractures appeared angular and consistent with overload. All flight control hardware not damaged by impact forces was found attached and secured.

The airspeed gauge indicated 103 knots, the rotor rpm gauge indicated 95 percent, the engine #1 gauge indicated 92 percent, and the engine #2 gauge indicated 90 percent.

The tail boom-Fenestron was located at the main wreckage site lying on its right side and oriented in a north-to-south direction. The Fenestron ring frame exhibited heat damage, and the fracture surfaces were consistent with overload.

The forward and aft ends of the composite-constructed tail rotor drive shaft were located within the main wreckage area. The forward two-thirds of the steel tail rotor drive shaft was also located within the main wreckage, while the aft one-third of the drive shaft was found about 120 feet east of the wreckage site.

System #1 and system #2 hydraulic pump/reservoir assemblies, along with all three main rotor actuators and the tail rotor actuator, were identified at the main wreckage site. No obvious preimpact mechanical anomalies were observed on any of the hydraulic system components.

The main transmission was fragmented, and all pieces were found near the main wreckage site. An examination of the main transmission components did not reveal any evidence of preimpact heat damage or internal scoring. All main transmission gear/component damage was consistent with impact and/or overload. The main transmission anti-resonance isolation system mounts were examined on site, with no evidence of preimpact damage observed. The damage to the main transmission torque strut was consistent with impact and overload.

The main rotor shaft and collector gear remained attached to the main rotor hub. All main rotor pitch links remained attached to the rotating swashplate and their respective main rotor blades.

The fuel system was destroyed by impact and the postcrash fire.

Both of the helicopte's engines were located within the main wreckage below the transmission deck. The left engine (#1 engine) had shifted inboard and somewhat below the right engine (#2 engine). Both engines exhibited damage from impact forces and the postcrash fire.

The #1 engine's front and rear mounts had separated at the airframe connections. The starter generator had separated from the engine at its mounting clamp, and the shaft exhibited torsional shearing. The oil and fuel filter by-pass indicator covers were broken. No indication of bypass was observed. The fuel valve assembly was intact and fire damaged. The compressor, gas generator turbine, and power turbine could not be rotated by hand. The blades of the compressor centrifugal wheel were intact. Foreign material was observed in the air intake casing and on the compressor blades. The power turbine containment shield was intact, and no blade shedding was observed. The turbine blades were discolored and contaminated with foreign material. The outlet diffuser and exhaust pipe were dented inboard at approximately the 11 o’clock position. The exhaust pipe was observed intact and elongated.

The #1 engine-to-aircraft transmission shaft was torsionally sheared at the rear flexible coupling. The forward connection had separated at the flange connection to the aircraft transmission. Examination of the main rotor rpm indicator in the cockpit revealed a reading of 91 percent for the #1 engine. The engine electronic control unit (EECU), part number (P/N) 70EMFO1160, S/N 8ALD0323CE, exhibited damage from impact forces and the postcrash fire. Subsequent to the completion of the onsite examinations, the engine was placed in a shipping container for shipment to the Turbomeca USA facility in Grand Prairie, Texas, where a more detailed examination was conducted on April 21, 2010.

The #2 engine’s right front engine mount had separated at the engine connection. The left front and rear engine mounts were separated at the airframe connection. The module one casing was separated at the mating flange to the module two, exposing the forward end of the power turbine shaft. The starter generator, hydro mechanical fuel unit (HMU), oil pump, and filter support block had separated with broken pieces of the module one casing. The starter intermediate gear and HMU drive gear were located in the wreckage below the engine. The oil and fuel filter bypass indicator covers were broken. No indication of bypass was observed. The fuel valve assembly was intact and fire damaged. The compressor, gas generator turbine, and power turbine could not be rotated by hand. The blades of the compressor centrifugal wheel were intact. Foreign material was observed in the air intake casing and on the compressor blades. The power turbine containment shield was intact, and no blade shedding was observed. The turbine blades were discolored and contaminated with foreign material. The outlet diffuser and exhaust pipe were crushed. The exhaust pipe extension had separated at the firewall.

The #2 engine-to-aircraft transmission shaft was torsionally sheared approximately 3 inches forward of the flexible coupling at the engine connection. The spline connection was on the engine power drive. The flexible coupling on the forward end of the shaft had separated, leaving the flange connected to the aircraft transmission. Examination of the main rotor rpm indicator in the cockpit revealed a reading of 92 percent for the #2 engine. The EECU (P/N 70EMFO1160, S/N 8ALD020CE) exhibited damage from impact forces and the postcrash fire. Subsequent to the completion of the onsite examinations, the engine was placed in a shipping container for shipment to the Turbomeca USA facility in Grand Prairie, Texas, where a more detailed examination was conducted on April 21, 2010.

Main Rotor Blades

All main rotor blade roots remained attached to the hub, and all main rotor blade bolts were intact. A section of the blue main rotor blade separated due to impact forces and came to rest at the main wreckage site. The three remaining main rotor blades (yellow, red, and green) remained attached to the hub but were missing some section of the respective blades, primarily at the outboard tip area.
Main Rotor Blade S/N 2472 (yellow blade)

A large portion of the blade exhibited impact damage and chordwise scoring. The outboard section of the blade separated about 14 feet 5 inches from the center of the blade’s attachment bushing. Pieces of the blade and blade tip were observed along the wreckage path. The blade’s cuff and dampers exhibited impact damage and contact marks.

Main Rotor Blade S/N 767 (green blade)

A large section of the blade exhibited impact damage and chordwise scoring. The blade was damaged/separated about 10 feet 6 inches outboard from the center of the blade attachment bushing. A 15-inch section of the blade was located next to the main wreckage, separated from the blade. The blade's cuff and dampers exhibited impact damage and contact marks.

Main Rotor Blade S/N 768 (blue blade)

The majority of the blue main rotor blade was observed separated from the hub and found next to the main wreckage site. The blade exhibited impact damage about 20 inches outboard from the center of the blade’s attachment bushing and separation about 55 inches from the center of the blade’s attachment bushing. A large portion of the blade was observed to have sustained impact damage and chordwise scoring. The outboard tip of the blade had separated about 15 feet 5 inches from the center of the blade’s attachment bushing. The blade tip was located with the wreckage debris during a February 17, 2012, follow-up examination of the helicopter. Granular pieces of the tungsten tip weight was observed and remained adhesively bonded to the composite structure of the blade near the tip. Pieces of the static and chordwise balancing chamber remained attached to the blade, but the trim weights were not observed. Aircraft wiring was embedded in the outboard spar near the tip separation. This was identified as the #1 hydraulic system wiring from the main transmission deck area. The blade’s cuff and dampers exhibited impact damage and contact marks.

Main Rotor Blade S/N 766 (red blade)

The entire length of the red blade was observed at the accident site. The blade’s trim tab was also observed; however, its static discharger was broken off. A large portion of the blade exhibited impact damage and chordwise scoring. The blade tip exhibited leading-edge impact damage and some apparent visible paint transfer. The blade's cuff and dampers were observed to have impact damage and contact marks respectively.

MEDICAL AND PATHOLOGICAL INFORMATION

Autopsies were performed on the pilot and the passengers at the Maricopa County Office of the Medical Examiner, Phoenix, Arizona, on February 17, 2010. The cause of death for each of the deceased was listed as "Blunt impact injuries."

The FAA's Civil Aeromedical Institute in Oklahoma City, Oklahoma, performed toxicology tests on the right front cockpit seat pilot-in-command and the non-rated male passenger who occupied the left front cockpit seat. The results of the testing for specimens received were negative for all tests conducted.

A review of medical records, dated July 28, 2009, which were supplied by the Asthma and Allergy Institute of Scottsdale, Arizona, revealed that the 5-year old female passenger’s weight was recorded as 39.5 pounds.

SURVIVAL FACTORS

An NTSB survival factors specialist reviewed the pilot seat range of motion; collective position, displacement, and force characteristics; cyclic position and displacement; restraint systems and restraint system use; and interior configuration.

The survival factors specialist's report noted that the range of motion of both pilot seats forward and aft is 4 inches horizontal. Additionally, the seats were able to move vertically a total of 2.3 inches. The report further stated that the collective is a part of the helicopter flight control system, which changes the pitch angle of all of the main rotor blades and results in an increase or decrease in power and main rotor lift. Each of the collectives is located on the left side of both cockpit seats. For the front left seat, the collective is located between the seat and the left door. During an examination of an exemplar helicopter, the spacing between the side of the seat and the left door was about 7 inches. The report also revealed that the collective friction and force characteristics for the push direction were 2.2 to 3.1 pounds of force and 5.0 pounds for maximum control force. Additionally, the displacement of the collective from the up position to the full-down position is about 9.5 inches. In the full-down position, the collective is at a 26 degree angle relative to horizontal, and in the full-up position, the collective is at 48 degrees.

The report noted that the cyclic is part of the helicopter flight control system and functions similar to the control stick on an aircraft, which changes the pitch of the aircraft and subsequently the angle of attack. The cyclic is located between the pilot’s legs in the center of each cockpit seat. (For more information, refer to Figure 5 of the Survival Factors Specialist’s Factual Report, which is located in the public docket for this accident.)

In a review of the helicopter's restraint systems, the survival factors specialist's report notes that the two cockpit seats were equipped with four-point restraint systems, while the four passenger seats (two rear facing and two forward facing) were equipped with three-point restraint systems with shoulder harnesses. The ranch foreman who assisted in loading the helicopter reported that the pilot was restrained with his lap belt and shoulder harness, but he was unable to see how the helicopter owner in the left front seat and the female child seated on his lap were restrained. He did note that on previous flights, the helicopter owner would strap the child on top of him. The ranch foreman added that he did not observe how the passengers seated in the rear cabin were restrained in the helicopter.

The report also notes that a preaccident photograph of the accident helicopter showed the interior configuration between the two aft-facing seats directly behind the pilot seats. The report indicates that while a direct measurement was not taken before the accident and measurements were not possible after the accident due to the extensive damage to the aircraft, the spacing between the upper partition and the ceiling was estimated to be about 5 inches.

TESTS AND RESEARCH

During the course of the investigation, various components of the helicopter were examined.

ENGINES

On April 21, 2010, under the supervision of representatives from the FAA’s Rotorcraft Directorate Office, Fort Worth, Texas, both engines were examined at the Turbomeca USA, Inc., Grand Prairie, Texas, facilities.

The #1 Engine S/N 30167 (left engine)

An initial assessment of the #1 engine determined that the engine was incapable of being run on the test cell. The HMU was deemed incapable of being run on the test bench due to severe impact forces and fire damage.

Disassembly consisted of removal of the exhaust diffuser, turbine ring, turbine cooling ring, power turbine wheel, and power turbine nozzle guide vane to facilitate examination of the trailing edge of the gas generator turbine blades. Examination of the power turbine nozzle guide vane revealed gray deposits on the vanes. The power turbine wheel exhibited deposits on the blade surfaces and blade tip rub. The examination also revealed that the backside of the turbine blades and thermocouple probes were coated with a yellow substance, a sample of which was retained by FAA representatives and sent to the NTSB Materials Laboratory, Washington, DC, for analysis. An NTSB chemist reported that the exact identity of the material examined could not be conclusively determined. (For more information, refer to NTSB Report No. 12-046, which is located in the public docket for this accident.)

The #2 Engine S/N 300291 (right engine)

An initial assessment of the #2 engine determined that the engine was incapable of being run on the test cell. The HMU was deemed incapable of being run on the test bench due to severe impact forces and fire damage.

Disassembly consisted of removal of the exhaust diffuser, turbine ring, turbine cooling ring, power turbine wheel, and power turbine nozzle guide vane to facilitate examination of the trailing edge of the gas generator turbine blades. The examination revealed the backside of the turbine blades were stained, and deposits on the inner surface of the turbine case were present. Examination of the power turbine components revealed deposits in the nozzle guide vane and rub marks on the inner surface of the turbine ring. The power turbine wheel exhibited blade tip rub.

Turbomeca reported that as a result of their examinations, the engines revealed signatures of engine rotation during the accident sequence.

Component examinations

On March 9, 2010, under the supervision of a German BFU representative, the helicopter’s clock, attitude indicator, altimeter, rotor tachometer, and airspeed indicator were examined at the BFU laboratory in Braunschweig, Germany. The BFU representative reported that the clock showed 1457 before and after opening of the case of the clock. There were no impact marks of the needles. The attitude indicator showed a bank of 157 degrees inverted. This indication was fixed, and it was not possible to rotate the inner roll. There were no impact marks of the needles.The altimeter: showed 2,000 feet; there were no impact marks of the needles. QNH (barometric pressure adjusted to sea level) was 29.97 In/Hg and 1015 hPa (Hectopascal pressure unit). Due to the destruction of the component, it was not possible to look for impact marks. The rpm indicator before opening the case indicated that "R" was 94.5 percent, "1" was 92 percent, and "2" was at 90 percent. After opening of the case, "R" was 94.5 percent, "1" was 91.5 percent, and "2" was 0 percent. Before opening the case, the airspeed indicator read 103 knots, and after opening the case, the indicator read 115 knots. There were no impact marks on the needles.

On March 24, 2010, under the supervision of a BFU representative and with Eurocopter Deutschland GmbH and Liebheer-Aerospace Lindenberg GmbH representatives present, the helicopter's two main hydraulic pressure supply pumps; single Fenestron actuator; and individual longitudinal, collective, and lateral hydraulic actuators were examined at the Liebherr-Aerospace facilities, Lindenberg, Germany. As reported by the BFU representative, no anomalies existed with any of the components that would have precluded normal operation. (For more information, refer to the BFU Laboratory Report, which is included in the public docket for this accident.)

Vehicle and Engine Multifunction Display (VEMD)

The helicopter was equipped with a Thales VEMD. The component was sent to the NTSB Vehicle Recorders Division, Washington, DC, for examination and download of information.

An NTSB vehicle recorders specialist reported that the VEMD is a multifunction screen installed on the instrument panel and is designed to manage essential and nonessential vehicle and engine data. Additionally, the VEMD is a dual-channel system, and each channel stores data and failure information on nonvolatile memory (NVM). The flight report, failure reports, and over-limit reports (on certain models) are stored in the VEMD.

The NTSB specialist reported that the VEMD was extensively damaged from impact forces. Due to the damage to the instrument panel, some components were not included in the shipment. The external case was crushed, and the circuit boards were cracked with multiple components missing.

One circuit card that contained NVM data was identified and inspected. The inspection revealed that the NVM chip was cracked along the top and right side. When the NVM chip was extracted from the circuit card, it was observed that the chip was cracked along the top and completely through the device. The chip was then X-rayed to determine the internal damage to the unit. The X-ray image did not show any visible damage to the die. Multiple bond wire connections were broken between the die and the leads. Multiple attempts were made to reconnect the broken bond wires using a probing station. However, due to the damage to the NVM chips, data extraction efforts have been unsuccessful, and at the time of this report no data has been recovered.

Warning Unit (WU) and Warning Unit Data

The helicopter was equipped with a WU, which consists of two fire warning emergency off switches and eight warning lights that illuminate red in case of activation. Each warning indication simultaneously initiates a gong audio signal. The ROTOR-RPM warning monitors a total of three limit values and reacts in various (visual and audio) ways depending on which limit value is exceeded or gone below.

All visual and audio warnings, up to 31 events, are recorded in the NVM of the WU until overwriting occurs. The LIMIT warning of the Flight Limit Indicator (FLI) is recorded as well. The WU stores data strictly in sequence of occurrence without any built-in timeline.

The helicopter’s WU (P/N 6300-00-00) was shipped to the BFU, who oversaw the data download. The readout of the data was subsequently performed by L’Hotellier of Antony, France. After being supplied with the downloaded data from the BFU, Eurocopter indicated the following to the NTSB:
• The WU recorded a normal starting sequence
• The warnings recorded were consistent with the accident sequence.
• Only one limit exceedance was recorded, which was attributed to an engine limit.
• When the helicopter impacted terrain during the crash sequence, the low rotor rpm warning was on. If it had not been on, the last recording would show an empty line.
• The mast moment system of the helicopter was inoperative, as the mast moment limit would have been exceeded several times during the complete crash sequence

(For more information, refer to the Eurocopter Analysis Warning Unit Data EC 135 S/N 0094 document, which is located in the public docket for this accident.)

The WU data revealed no indications that a LIMIT warning for the mast moment exceedance was ever activated. SGA reported that the mast moment system had been inoperative at the time of the accident flight.

EECU

On June 8, 2010, under the supervision of an NTSB aerospace engineer, an examination of the helicopter's #1 engine (left) EECU, model EMC-35B (S/N 8ALD0323CE), was performed at the Goodrich Pump & Engine Control Systems facilities, West Hartford, Connecticut. A Turbomeca USA accident investigator and a Goodrich Pump & Engine Control Systems product support engineer attended the examination. The objective of the examination was to examine the condition of the memory integrated circuit (IC), which was located on the computer circuit card assembly (CAA), to determine if pertinent data stored on the IC could be extracted from the unit.

An initial external examination of the EECU revealed that it exhibited extensive structural damage, with significant thermal stress damage caused by the postcrash fire. Further examination revealed that after the CCA was exposed, the unit also exhibited significant structural and fire damage. The damage to the CCA was severe. The memory IC was then examined and found to be severely damaged as well, making the interrogation for any stored information not possible.

The helicopter’s #2 engine (right) EECU, model EMC-35B (S/N 8ALD0202CE), was severely destroyed by impact forces and the postcrash fire. The condition of the unit precluded any follow up examination.

NTSB Materials Laboratory Examinations

Subsequent to the helicopter wreckage being examined at the accident site, as well as in a follow up examination at a facility where it was stored in Phoenix, Arizona, various component parts were retained by the NTSB and shipped to the American Eurocopter facility in Grand Prairie, Texas, for further examination by an NTSB materials research engineer. As a result of this examination, which took place between May 11 and May 13, 2010, the engineer had selected parts sent to the NTSB Materials Laboratory, Washington, DC, for further examination.

At the NTSB Materials Laboratory, a detailed examination was conducted of the following helicopter components: (1) the aft end of the steel section of the tail rotor drive shaft, including the aft flexible coupling; (2) the forward end of the steel section of the tail rotor drive shaft, including the forward flexible coupling and forward bearing; (3) the left vertical endplate of the horizontal stabilizer; (4) tip pieces of the four main rotor blades; and (5) tip pieces of the four main rotor blades.

Tail Rotor Drive Shaft

The aft piece of the steel section of the tail rotor drive shaft was separated and located 75 feet east-northeast of the main wreckage. The thin metal leaves of the flexible coupling at the aft end were fractured in overstress at numerous locations, which would have separated the steel section of the drive shaft from the aft composite section of the shaft. A metal fragment, which was subsequently identified as a part of the nickel-cobalt alloy leading-edge protection strip from the yellow blade, was lodged in the flexible coupling. The metal fragment had been wedged into the coupling by an impact from the yellow blade as it swept from left to right over the tail boom.

The aft piece of the steel section of the tail rotor drive shaft was fractured about 28 inches forward of the flexible coupling. The fracture surface had a matte texture, which is consistent with overstress separation. The examination revealed that the deformation and fracture was consistent with a fracture occurring under torsion caused by rotation of the tail rotor drive shaft.

The forward end of the steel section of the tail rotor drive shaft where it was joined to the forward composite section of the tail rotor drive shaft was found with the main wreckage. One bolt (of six) was missing from the flexible coupling at the forward end of the steel section of the tail rotor drive shaft. The aft side of the hole for the missing bolt was elongated toward the center of the shaft and clockwise looking forward, with a lip pushed up at the elongation, consistent with indentation by the bolt at ground impact. A steel bushing that clamped the leaves of the flexible coupling at that location was also missing.

It was observed at the crash site that the #1 tail rotor drive shaft bearing remained loosely in position at the forward end of the steel section of the drive shaft, but the outer bearing race was separated from the ring frame structure at the forward end of the tail boom where it had been mounted. The outer race, inner race, cage, and balls were still assembled, and there was no apparent mechanical damage to the bearing, but the bearing surfaces were oxidized and the polymeric portions of the seals were missing as a result of the postcrash fire. The bearing could only be turned a small amount without a large force. The two bolts that had attached the bearing to the ring frame were not present. The two bolt holes in the attachment flanges (which are part of the outer race of the bearing) showed no elongation or other markings. The outer race of the bearing was significantly harder than the titanium bolts used to attach the bearing to the ring frame.

Left Vertical Endplate of the Horizontal Stabilizer

An examination of the endplate, which was constructed of woven carbon-fiber composite around a honeycomb core, was observed to have had three parallel cuts that passed through the top of the endplate and roughly in line with the top of the tail boom and the tail rotor drive shaft. The cuts were consistent with the tip of one or more main rotor blades passing through the endplate. The three cuts were about 10 to 12 inches long, with the top two cuts separated by 2.25 inches and the third cut 1.25 inches farther down. A fourth cut existed farther down near the level where the endplate was attached to the horizontal stabilizer. The fourth cut was approximately 14 inches long, as it was oriented at a different angle than the three cuts above it.

Main Rotor Blades

The helicopter had four main rotor blades, which extended 16.7 feet (200.8 inches) from the center of the mast. Portions of the four main rotor blades were examined, which included the outermost tip portions recovered for each blade. Many pieces of the tip of the yellow blade were recovered along the flightpath of the helicopter leading up the main crash site. All of the blade pieces from the green, blue, and red blades were recovered from the main crash site.

The green and blue blades were the most severely damaged. The red and yellow blades were more nearly intact. The tip pieces from the green, blue, and red blades had numerous chordwise scratches and impacts on the leading-edge protection strips, indicating that the main rotor blades were rotating at ground impact.

No indications existed of preexisting damage such as fatigue cracking on any of the main rotor blade pieces examined. All of the fractures in the leading-edge protection strips occurred along slant planes consistent with overstress separation.

The inner edges of all of the blade cuffs had marks consistent with contact with the corners of the mounting bolts, which were similar to contact marks previously seen by Eurocopter that had prompted a modification of the cuffs. The similarity of the marks between blades and their similarity to previously observed marks indicated that they likely occurred in flight. The inner edge of the yellow blade cuff had more severe impact marks (discussed below) that were unique and likely occurred at ground impact.

Yellow Blade (S/N 2472)

The NTSB materials engineer reported that the examination of the yellow blade (S/N 2472) revealed that the blade was generally intact from the blade root to a jagged fracture some 20 to 50 inches from the tip. The tip of the blade was missing at the main crash site, but 10 pieces from the tip were recovered along the flightpath leading up to the crash site and matched to the tip area of the yellow blade by aligning their fracture surfaces. All of the main rotor blades pieces recovered along the flightpath were from the yellow blade.

As noted above, a fragment of nickel-cobalt alloy main rotor blade leading-edge protection strip was recovered from the flexible coupling attaching two portions of the tail rotor drive shaft at the transition between the tail boom and the Fenestron. Aligning the fracture surfaces showed that this fragment mated to the yellow blade tip pieces recovered along the flightpath leading to the crash site.

The yellow blade cuff was fractured just outboard of the pitch control linkage, and the internal cruciform beam in the vicinity of the fracture was delaminated. On the cuff, the fracture surfaces were all relatively flat, with few extended fibrous regions, consistent with fracture primarily under local bending as a result of axial compression of the cuff. The inboard edge of the cuff retained several sets of deep indentation marks indicating severe contact with the main rotor hub, again consistent with compression along the axis of the blade. None of the cuffs from the other blades showed similar severe impact marks in this location. The interior cruciform beam was not fractured, but the unidirectional glass-fiber composite material was severely delaminated in the vicinity of the fracture in the cuff, also consistent with compression along the axis of the blade.

Green Blade (S/N 767)

The green blade was generally intact over approximately half its length (from the root to a point between 100 inches and 124 inches outboard). The blade was fractured and fragmented between that location and the inboard end of the 17-inch-long tip piece of the green blade, which was recovered at the crash site.

Blue Blade (S/N 768)

The blue blade was badly damaged about 28 inches from the root, with a fracture in the blade cuff and bending damage in the interior cruciform beam. The blade was also fractured and separated about 64 inches from the root. Outboard of that fracture, the leading edge portion of the blue blade was generally intact out to a position about 190 inches from the root. Behind the leading edge, the blade was fragmented, with much of the trailing-edge portion of the blade not recovered or not identified.

While the blue blade tip was not located during the initial stages of the investigation or during follow-up searches of the crash site and the surrounding area, the blade tip was subsequently discovered on February 17, 2012, among the stored wreckage at a salvage facility in Phoenix, Arizona. The recovered blue blade tip piece was about 15 inches long and exhibited severe impact damage at the leading edge.

Red Blade (S/N 766)

At the crash site, the red blade was observed to extend continuously from the blade root to the tip. The leading-edge protection strip was fractured and indented consistent with an impact on the leading edge at about 189 inches from the root. There was another indentation consistent with impact at the leading edge further inboard, about 179 inches from the root.

Examination for Paint Transfer

Pieces of the left vertical endplate from the horizontal stabilizer and tip pieces from all four main rotor blades were sent to the Federal Bureau of Investigation (FBI) laboratory to determine whether any paint from the endplate had been transferred to the main rotor blades. No paint transfer was identified, but some fragments of carbon fiber composite material were found on some of the yellow blade pieces. It was noted that the endplate was constructed of carbon-fiber composite materials, but there was also a strip of (conductive) carbon-fiber composite material running spanwise along each blade for lightning protection. A sample of the lightning protection strip was also provided to the FBI laboratory. These two carbon-fiber composite materials could not be discriminated, so the source of the carbon-fiber composite fragments could not be conclusively identified. (For more information, refer to the FBI laboratory report, which is attached as Appendix A of the Materials Laboratory Factual Report No. 11-014, which is located in the public docket for this accident.)

Eurocopter Study: Main Rotor Disc Divergence from Normal Plane of Rotation

During the investigation, Eurocopter conducted a study to determine what condition or event could have caused the main rotor disc to divert from its normal plane of rotation and subsequently strike the endplates and the tail rotor drive shaft. The study started with an assumption that contact between the main rotor blades and the endplates or tail boom could only be possible when a pilot carries out an extreme aft cyclic input.

Eurocopter initially studied the possibility of an occurrence termed as an "Aggressive Pull Aft" maneuver during fast forward flight, where the pilot aggressively pulls the cyclic pitch from a forward position fully to the aft stop.

In a simulation model, it was demonstrated that during an "Aggressive Pull Aft" maneuver alone, sufficient clearance to the endplates and tail boom existed. As a result, Eurocopter concluded that another flight maneuver must precede the "Aggressive Pull Aft" in a sequence to explain the excessive flapping required that would result in the main rotor blade collision with the endplates and the tail boom.

To address this issue and through approved engineering and simulation models, Eurocopter provided the investigation team with two accident flight maneuver sequence scenarios. In each case, the maneuver sequence starts at the same cruise flight condition (fast, forward flight, high collective pitch, and forward longitudinal cyclic):

Scenario #1: Push cyclic full remaining range forward and then pull full range of longitudinal cyclic control backward (cyclic push/pull maneuver).

Eurocopter determined that the control sequence in scenario #1 resulted in an increased blade deflection when compared to the "Aggressive Pull Aft" maneuver but still showed enough clearance between the main rotor disc and the tail section of the helicopter.

Scenario #2: Sudden lowering of the collective to near the lower stop, followed by a simultaneous reaction of nearly full up collective and nearly full-aft cyclic (longitudinal).

Eurocopter determined that the control sequence in scenario #2, based on known conditions at the time of the accident and the laws of physics and aerodynamic principles, was the only maneuver that would result in main rotor blade contact with the endplates and tail boom.

(For more information, refer to the Eurocopter submission, pages 27 thru 45, which is located in the public docket for this accident.)

Members of the investigation team, which included the NTSB chief scientist, an NTSB structures engineer, an NTSB materials engineer, a BFU accredited representative, and a representative from the FAA's Rotorcraft Directorate, conducted a thorough review of Eurocopter's study and concurred with the findings of the study as presented.

Weight and Balance Study

An NTSB air safety investigator conducted a weight and balance study of the accident helicopter. On the day of the accident flight, it departed with an unknown quantity of fuel. However, based on historical fueling records supplied by SGA personnel, the investigator formulated three scenarios, each of which would be representative of the weight and balance condition of the helicopter when it arrived at the accident site. The investigator reported that the helicopter's gross weight in all three scenarios differed only by about 100 pounds and that all weights fell within the flight envelope and below the helicopter’s maximum certified gross weight at the time of the accident. (For more information, refer to the Weight and Balance Study, which is located in the public docket for this accident.)

Immersive Witness Interview (IWI) Study

The IWI technology was developed to approximate the observed flightpath of aircraft based on eyewitness observations. The technology was developed in cooperation with the German Federal Armed Forces Flight Safety Division, with the support of the German Air Force and Eurocopter. The IWI technology uses witness reports, which are transferred into a 3-D environment to calculate situation reconstruction by using lines of sight and to validate witness observations.

About 14 months after the accident, five witnesses participated in the IWI process. At the time, Eurocopter's primary objective was to demonstrate the IWI technology to the NTSB. In summary, Eurocopter provided the following relative to the interviews conducted on April 19th and April 20th, 2011:

In most cases, the "pop pop" sound drew the witnesses' attention to the helicopter. The investigation revealed that the "pop pop" sound occurred at some point around 22:03:23. Regarding to the longitudinal and lateral position of the helicopter, the witness observations were generally consistent with the radar data. Witness #4 was generally below the helicopter and began to move to a higher elevation, looking up to observe the helicopter. This is a challenging perspective for a witness from the standpoint that the witness had limited reference objects, which changed over time as he moved. Witness #5 reported the altitude of the helicopter"s flight before the accident sequence to be generally lower than what was recorded by radar. As it would have taken the witness several seconds after the "pop pop" sound to move to his vantage point, it is unclear at which point during the flightpath he observed the helicopter. Based on the observations of witnesses #1 and #2, the helicopter’s flightpath at the end was almost straight down.
(For more information, refer to the Immersive Witness Interview Study, which is located in the public docket for this accident.)

ADDITIONAL INFORMATION

Responsibility and Authority of the Pilot-in-Command

Title 14 CFR 91.3: Responsibility and authority of the pilot in command.
The pilot in command of an aircraft is directly responsible for, and is the final authority as to, the operation of that aircraft.

According to Federal Aviation Administration Advisory Circular 60-22, Aeronautical Decision Making (ADM), ADM is defined as systematic approach to the mental process used by aircraft pilots to consistently determine the best course of action in response to a given set of circumstances. Additionally, the term Poor Judgment Chain is defined as a series of mistakes that my lead to an accident or incident. Generally associated with this chain are two basic principles: (1) one bad decision often leads to another, and (2) as a string of bad decisions grows, it reduces the number of subsequent alternatives for continued safe flight.

AC 60-22 also notes that there are a number of classic behavioral traps into which pilots have been known to fall. Pilots, particularly those with considerable experience, as a rule always try to complete a flight as planned, please passengers, and meet schedules. These tendencies ultimately may lead to practices that are dangerous and often illegal, and may lead to a mishap.

According the FAA Risk Management Handbook, (FAA-H-8380-2), "Poor decision making by pilots has been identified as a major factor in many aviation accidents." It further states, "External pressures are influences external to the flight that create a sense of pressure to complete the flight, such as a passenger the pilot does not want to disappoint. Management of external pressure is the single most important key to risk management because it is the one risk factor category that can cause a pilot to ignore all other risk factors."

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