On June 16, 2010, about 1440 mountain daylight time, a Kaman K-1200 helicopter, N134WC, impacted the terrain about five miles west of Donnelly, Idaho. The commercial pilot, who was the sole occupant, was killed in the accident sequence, and the helicopter, which was owned and operated by Woody Contracting Inc., sustained substantial damage. The 14 Code of Federal Regulations Part 133 long-line logging flight had been airborne for about an hour and 40 minutes. The flight was taking place in visual meteorological conditions. No flight plan had been filed.

According to the four witnesses near the site, the helicopter, which was using a 200 foot long-line, had just lifted a large log off the ground when the accident sequence began. Two of the witnesses were looking directly at the helicopter as it lifted the log off the ground, and the other two looked immediately toward the helicopter after hearing a very loud noise that coincided with the beginning of the accident sequence. The pilot of the helicopter, who intended to take the log to a collection landing area about one-quarter mile away, had already attempted to lift the log off the ground once, but other logs interfered with the lift, so he set it back down. The pilot then asked the hooker to reset the choker nearer to the end of the log, which according to one of the witnesses would cause the log to hang more straight down during the lift. The hooker therefore reset the choker nearer to the butt end of the log, whereupon the pilot transmitted that, "That is more like it." Soon thereafter, the pilot again began to lift the log. According to the two witnesses that had a good view of the log itself, the accident sequence began just after the log was completely off the ground.

The two witnesses who were looking directly at the helicopter perceived the accident sequence to have begun when they saw what appeared to be an outboard section of one of the rotor blades separate and fly through the air. This was followed a "split second" later by a very loud bang or pop. To the two witnesses who were not looking directly at the helicopter, the sequence began when they heard the very loud noise. Both of these individuals, one of whom was about one-quarter mile away, and the other (the hooker) who was within 50 feet of the helicopter, immediately turned to look directly at the helicopter. The first thing that both of these individuals saw was a piece (pieces according to the hooker) of the rotor blade flying through the air. Almost immediately after the separation of the rotor blade piece (pieces), other portions of the helicopter's blades and structure began to separate. The helicopter then went out of control, descended toward the terrain, ultimately coming to rest inverted. There was no fire.

According to all four witnesses, there were no unusual engine or rotor sounds prior to the initiation of the accident sequence. They all stated that everything looked and appeared normal prior to the separation of the first rotor blade piece. Two of the witnesses made the point of saying that they had worked around helicopters for a long time, and felt that they would have recognized any abnormal change in engine or rotor sound. One witness made the point that he had heard unusual engine and rotor sounds from other helicopters over the years, but that this time everything sounded perfectly normal.

Of the two witnesses who were looking directly at the helicopter, the one closest to the event was asked if it appeared that the pilot ascended faster than normal or tried to jerk the log off the ground. He responded that the pilot had not done either of those things, and that the lift looked like any other lift involving a single large log. The witnesses that were in position to judge how close the helicopter was to any trees were adamant that its rotor blades had not come in contact with a tree or any other obstacle.


The pilot, a 64 year old male, held a Commercial pilot certificate, with a helicopter rating and an airplane single engine land rating. His last FAA airman's medical, a class 2, was performed on 3/18/2001, and was issued without limitations or waivers. His total pilot time as of the date of the accident was estimated to be 27,600 hours, with 12,600 hours in the Kaman 1200. His last flight review was on December 8, 2008, in a Robinson R-22.


The aircraft was a 1996 Kaman Aerospace K-1200 KMAX helicopter, serial number A94-0006, registered as N134WC, and powered by a Honeywell/Lycoming T53-17A1 engine, rated at 1,500 horsepower. The helicopter's last progressive inspection under its Approved Airworthiness Inspection Program (AAIP) was on June 15, 2010. At the time of the accident, the airframe had accumulated 24,053 hours.


At the time of the accident, the helicopter pilot was working under a 12,000 foot overcast ceiling, in winds that varied from calm to variable at six knots. Visibility was more than ten miles. The temperature was about 45 degrees F.


The helicopter lifted the log from the ground about the 6,300 foot level of the southeast facing down-slope of North Business Mountain, at about 44 degrees, 45.771 minutes north by 116 degrees, 11.138 minutes west. At the termination of the accident sequence, the airframe came to rest inverted over the top of the log it had lifted, about 410 feet southeast of the lift location, at an altitude of about 6,180 feet. The airframe wreckage included the main fuselage, the nose and cockpit area, the main and nose landing gear, the engine, transmission, and right rotor mast pylon, both left and right stabilizers, and the tail boom, including the vertical fin and rudder.

Both horizontal stabilizers were fully attached, and both vertical stabilizers were fully attached to their respective horizontal stabilizers. Both horizontal stabilizers had rotated about 180 degrees so that the leading edge of both were facing aft, and the bottom tip of both vertical stabilizers were facing toward the top of the fuselage. The outboard two-thirds of the trailing edge of the right horizontal stabilizer was bent down to about 90 degrees, and trailing edge of the top half of the associated vertical stabilizer was bent outward to about 90 degrees. The top half of the right vertical stabilizer was bent inward toward the fuselage at about 45 degrees, but the bottom half was in its normal alignment. The left horizontal stabilizer was undamaged, and its associated vertical stabilizer was undamaged except for an area near the forward portion of the bottom tip where it had been crushed inward about 10 inches by impact with a limb of the log that was being lifted. The tail boom had twisted counter-clockwise (looking forward) just aft of the tail boom mounting points, and its structural skin had been almost entirely ripped apart at that location. All five tail boom mounting bolts were connected through their fittings, and the fittings themselves were undamaged. The remainder of the tail boom, aft of where it had twisted/torn, was undamaged, and the rudder was fully connected and free to pivot. There was no evidence of any flight control surface anomaly or malfunction.

The belly hook was in the open position, and the bow shackle on the upper end of the long-line was released from the hook. The long-line itself was wrapped around the fuselage and left stabilizer, then strung up into some nearby birch trees, and then back down to the remote hook. The remote hook was connected to the choker, which was still attached to the log about three feet from its butt end. There were no marks or damage on the long-line that would have been consistent with it coming in contact with the rotor blades.

The transmission was in its correct location, but the KAFlex coupling between the engine and transmission had sheared, with the associated adapter rings still being attached to the transmission and engine respectively. The forward engine mounts had separated from the helicopter's airframe structure, and the aft engine mounts had been torn from the engine diffuser housing. The right pylon was attached to the transmission, but was partially separated along the outboard edge of the transmission top cover and pylon base. The mast still protruded from the top of the pylon, and was imbedded in the dirt under the inverted fuselage. The rotor hub assembly had fractured releasing blades 94A and 94B. The majority of the structure of blades 94A and 94B was found southwest of the fuselage about 470 feet and 440 feet respectively. The most outboard six and one-half feet of blade 94B was located at a later date about two-tenths of a mile south of the fuselage.

The left pylon, including the upper cover and the planet gears, had separated from the transmission where the pylon bolts to the transmission top cover, and was located about 190 feet east of the fuselage section. It had come to rest inverted (mast tip imbedded in the soil), with about 12.5 feet of blade 169B and 3.5 feet of blade 169A still attached. The damper tube black end was attached to blade 169B, and the fractured damper tube white rod end was attached to blade 169A. The U-crank was still attached to blade 169B, but the U-crank on blade 169A had broken off and was lying alongside the blade section, connected only by the idler assembly. The grips and teeter pins associated with both blades were intact. The majority of the remainder of blade 169A was located in an area about 250 feet west of the fuselage. The remainder of blade 169B was more widely scattered, with much of the blade not recovered, including any structure that could be positively identified as being outboard of station 260.

The servo flap for blade 94A was attached to both of its mounting brackets, and both brackets remained securely attached to the blade afterbody. The leading edge/spar of the servo flap for blade 169A was attached to both of its mounting brackets, and both brackets were securely attached to the blade afterbody. The servo flap afterbody for blade 169A had separated along a straight span-wise line about one inch aft of its pivot point along its top skin, and about two inches aft of the pivot point along its lower skin. The servo flap afterbody itself was found lying near the tail of the helicopter. Most of the servo flap for blade 169B was found about 230 feet north of the fuselage. Its leading edge/spar, including the stainless steel protection strip, was all present, and it retained about two-thirds of the outboard portion of its afterbody. The inboard end of the leading edge/spar itself was attached to the inboard mounting bracket and pitch change linkage, but it had fractured about four inches outboard of the pivot bearing, and was retained only by the pivot rod. Its outboard pivot bearing had been pulled from the associated mounting bracket, and the mounting bracket itself was found securely attached to about an 18 inch section of blade afterbody. The servo flap for blade 94B was fragmented and recovered in pieces. Its inboard fitting was securely attached to the blade body, but the inboard end of the flap leading edge/spar had separated at the inboard pivot bearing, and was retained only by the pivot rod and the pitch change linkage. The leading edge/spar protection strip was all present, although deformed and bent forward along its outboard most eight inches. The leading edge/spar structure was present, except for about its most outboard four inches. The outboard hinge fitting was recovered, but it was completely detached from the servo flap pivot bearing. Its forward one-half was attached only to a piece of blade honeycomb structure barely wider than the fitting itself, and its forward most three inches was bent up and aft. The leading edge of the fitting and its forward attachment bolt and nut were damaged and deformed toward the trailing edge, in a manner consistent with impact from a blade rotating counter to it.

After the on-scene investigation was completed, the wreckage was recovered to the facilities of SP Aircraft, in Boise, Idaho, where the NTSB Investigator-In-Charge (IIC), oversaw a layout inspection of the entire wreckage, including the engine. At the completion of that inspection, blades 94B and 169B, which had witness marks indicating that their leading edges had come in contact with each other, were shipped to the Kaman Aerospace Corporation, in Bloomfield, Connecticut. There they underwent a series of X-ray inspections, overseen by a representative from the Federal Aviation Administration (FAA) Manufacturing Inspection District Office at Windsor Locks, Connecticut (MIDO -41). The blades were subjected to the X-ray process in order to determine whether there were any preexisting anomalies within the interior of the blades, and to specifically detect any existing span-wise/chord-wise cracks or span-wise delamination in the laminated spruce spar. At the completion of the X-ray inspection, blades 94B and 169B were shipped to the NTSB's Materials Laboratory in Washington, D.C., for further examination and analysis. Upon completion of the Materials Laboratory examination process, blades 94B and 169B were returned to the facilities of SP Aircraft, where the investigative team performed another inspection of the major components of the helicopter's drive train, flight control system, and rotor blades. That examination was overseen and directed by an NTSB Materials Research Engineer. In addition, four panel-mounted instruments were removed from the wreckage and sent to the facilities of Howell Instruments, Inc., in Fort Worth, Texas, for inspection and attempted data retrieval. That activity was overseen by an NTSB Air Safety Investigator from the NTSB's South Central Region.

The aforementioned series of inspections and examinations determined the following:
1. All fractures associated with the separation of the left pylon from the transmission were consistent with overstress, with no indication of preexisting fatigue cracking found.
2. Both the forward and aft left pylon supports failed in overstress, with no indication of preexisting fatigue cracking found.
3. The pylon-to-pylon shear tie failed in overstress on its left side and on the upper channel of its right side. Its right side lower channel cracked, but did not fully fracture. There was no evidence of preexisting fatigue cracks.
4. Turning input to the transmission freely turned the right main rotor shaft and the right azimuth assembly.
5. Both the azimuth assembly and the five planet gears at the bottom of the left pylon turned freely.
6. None of the components of the left side planetary gear system exhibited any visible damage to suggest misalignment, binding or excessive wear.
7. The left and right transmission chip detectors were free of any contamination or metallic particles.
8. All the pivots, bellcranks, rod ends, and bearings associated with the cyclic and collective controls and the horizontal stabilizer control systems were still connected, except where separated by overstress fractures. The aforementioned components all moved freely except where binding occurred due to impact damage.
9. All four main rotor blades fractured in multiple locations, with the servo flap push-pull tubes generally fracturing adjacent to the rod end at the inboard ends of the blades, where they were connected to the U-crank and idler assemblies. All the fractures in the servo flap push-pull tubes were consistent with overstress separation. Blades 94A and 94B separated from the mast assembly at, or just inboard of, the lead-lag pivot. Blades 169A and 169B separated from the blade root about two feet outboard of the end of the blade grip.
10. Both blade 169A and 94A had fractured in two similar locations (about station 50 to 60, and again about station 230 to 240), and neither blade showed evidence of coming in contact with another blade or airframe structure outboard of about station 65. The fractures at both locations on both of the blades were consistent with separation in overstress, and no evidence of a pre-crash anomaly was detected at any of the fractures. Inboard of Station 60, both blades displayed evidence of contact with one another. The bottom skin of blade 94A retained a 1 to 2 inch-wide curved scrape near blade station 52, which was consistent with contact with hardware on the left main rotor. Also, a fragment of glass-fiber composite skin and honeycomb that was wedged between the grip and the top of blade 169A was matched to a location on the lower skin of blade 94A between blade stations 60 and 65. In addition, on blade 169A there was damage to the U-crank and U-crank pivot bolt hole, along with other contact witness marks consistent with blade 94A impacting the U-crank of blade 169A, and pushing the U-crank toward the center of the mast.
11. Blade 94B separated from the right mast by overstress fractures in the hub assembly, and the blade itself fractured in two locations. The most inboard blade fracture, which revealed surfaces consistent with an overstress failure, was approximately at the same location as the inboard fractures on blades 169A and 94A (station 55). The outboard end of the blade section inboard of the fracture showed severe leading edge contact damage, and the bottom surface of the blade area just outboard of the fracture contained curving scrape marks consistent with contact with hardware on the left main rotor. The second fracture of the wood spar, at station 193, was associated with the blade body being fragmented into a number of pieces. The fracture in the spar itself at station 193 was consistent with an overstress tension failure near the leading edge, and consistent with an overstress compression failure near the trailing edge. The stainless steel leading edge protection strip, which was still attached to a piece of the blade consisting of station 239 to station 289 (the blade tip), was fractured at station 210, with deformation at the trailing edge of the strip indicating fracture under forward bending. The leading edge protection strip was distorted consistent with impact between stations 185 and 220 (on both sides of the fracture), with the most severe deformation adjacent to the fracture at station 210. The lower surface of the protection strip exhibited a smoothly curving witness mark initiating at station 243 and extending inboard to station 185, consistent with contact with blade 169B.
12. Blade 169B, inboard of a transverse fracture running from station 117 to about station 152, was still attached to the left mast/hub assembly. Outboard of the transverse fracture, blade 169B was the most severely fragmented, and much of the blade was either not recovered or not able to be identified. The bottom surface of the blade inboard of the transverse fracture did not show any evidence of contact marks that would have indicated that it had come in contact with hardware of the right main rotor system. At station 117, the wood spar fracture surfaces were jagged, consistent with tension overstress, toward the leading edge and lower surface, and were flatter, consistent with compression overstress, toward the upper trailing edge. A portion of the stainless steel blade leading edge protection strip, running from station 180.5 to station 225.5, was recovered (in two pieces). The strip, which was indented and deformed in a manner consistent with impact with the leading edge of blade 94B, also retained a vertical indentation 0.75 inch wide and 2 inches high at station 198, consistent with impact from the outboard 0.675-inch-wide servo flap hinge fitting on blade 94B. No identifiable pieces of blade 169B outboard of station 260 were recovered.
13. The X-ray inspection process did not reveal any evidence of pre-existing blade spar cracks or delaminations, nor did it detect any internal blade anomalies not directly associated with impact damage.
14. Inspection of the engine revealed rotational and thermal damage consistent with the production of power both at the initiation of and during the accident sequence. Examination of the engine inlet revealed extensive foreign object damage to the axial compressor and inlet guide vanes, and a number of airframe fittings and hardware pieces were found lying in the inlet against the remains of the inlet guide vanes. The first stage compressor blades were all missing, along with the first stage stator vanes. The second and third stage compressor blades showed heavy bending opposite the direction of rotation and the remaining stator vanes were bent in the direction of rotation. The power turbine, reduction gearbox, and output shaft were free to rotate, and all rotated smoothly by turning the remains of the KAflex coupling and output shaft of the engine. The power turbine was intact, and had fine light colored debris adhering to the surfaces in the area of the 2nd stage power turbine wheel. There was also a spray pattern consistent with melted metal that had adhered to the 2nd stage nozzle, and a silver-gray colored material had collected on the inside of the blade tip shrouds. The power turbine governor drive shaft was removed, and found to be intact and in good condition. Continuity was confirmed between the power turbine group and the gearbox, and to the power turbine governor, by rotating the engine output shaft and witnessing the rotation of the splined gearshaft.
15. Of the four panel-mounted instruments that were removed for inspection and attempted data retrieval (H1900K-22 EGT Indicator, H1901-1 NG RPM Indicator, H1973-1 Load Indicator, H1943-1 Torque Indicator), all had sustained varying degrees of impact damage, up to and including multiple fractures of printed circuit boards. According to a representative of Howell Instruments, no recorded data points were able to be recovered from the EGT Indicator, the Load Indicator, or the Torque Indicator. A Peak Engine Speed (NG) of 101.4 percent was recovered from the NG RPM Indicator, but, according to Howell Instruments, that data point could not positively be determined to be associated with the accident sequence. A Maximum Load Warning Limit of 2,724 kilograms (5,992.8 pounds) was recovered from the Load Indicator, but that was a data point preset by software. Exceedance flags were triggered on both the Load Indicator and the Torque Indicator, but there were no ball codes recorded to further define the exceedance event, and according to Howell Instruments, it cannot be determined if the flags were associated with the accident sequence, some other sequence, or the result of corrupted data.


An autopsy was performed under the authority of the Valley County Coroner's Office, with the cause of death being blunt force trauma, and the manner of death being accidental.

The FAA's Civil Aerospace Medical Institute (CAMI) performed a forensic toxicology examination on specimens retrieved from the pilot. The results of that examination were negative for carbon monoxide and cyanide in the blood, negative for ethanol in the vitreous, and negative for listed drugs in the urine.



At the request of the NTSB IIC, the operator weighed the 200-foot long-line, the 30 foot choker, and the log that was being lifted. According to the operator, the long-line and associated hardware weighed 146 pounds, the choker weighed 15 pounds, and the log itself weighed 6,330 pounds. Based on these weights, the external load being lifted was 6,491 pounds. According to the helicopter's manufacturer and the limitation printed on the side of the helicopter, the maximum permissible external load was 6,000 pounds.

According to the calculations of the operator, the helicopter itself weighed 5,797 pounds at the time of the accident. With the addition of the 6,491 pound total weight of the external load, the helicopter's estimated total gross weight at the time of the accident was 12,288 pounds, which is 288 pounds over maximum allowable gross weight.


K-Max rotor blades are put into service as matched pairs. Some blades have been re-matched when their original opposite blade was removed from service. When that was done, the new pairing was re-designated as a new serial number pair, and the time-since-new assigned to that pairing going forward was the time-since-new of the higher time of the two blades.

The blade pairings on N134WC were serial numbers 94A/94B, and serial numbers 169A/169B. Blade pair 94A/94B was an original pairing with a zero time installation date of 15 February 1998. Blade pair 169A/169B was a re-matched pair. Blade 169A was originally serial number 65A, with an initial zero time installation date of 1 December 1996. Blade 169B was previously serial number 145B, and originally had been blade serial number 47B, with an initial zero time installation date of 26 July 1995. Records indicated that at the time of the accident, blade pair 94A/94B had accumulated 4,803.2 hours since new, and 1,226.5 hours since overhaul. The records indicated that blade pair 169A/169B had accumulated 9,633.2 hours since new, and 256.1 hours since overhaul. According to records provided by Kaman, at the time of the accident, blade pair 169A/169B had accumulated the eighth highest time since new of the known operational inventory. The highest time blade pair (72A/72B) had accumulated 11,225.0 hours since new (about 1,592 more hours than blade pair 169A/169B).

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