N775AR

HISTORY OF FLIGHT

On July 16, 2013, about 0955 mountain daylight time, a Tamarack UH-1H helicopter, N775AR, was substantially damaged after a loss of control and ground impact near Dove Creek, Colorado. The pilot, the sole occupant, was fatally injured. The helicopter was registered to BVDS Incorporated and operated by Billings Flying Service under the provisions of 14 Code of Federal Regulations Part 133 during a seismic survey operation. Visual meteorological conditions prevailed for the local flight, which departed without a flight plan from a staging area near Dove Creek, Colorado, about 0953.

According to ground witnesses at the survey location, the pilot flew inbound to their location from the left seat, hoisting a basket load with a long line rope. The pilot overshot the intended drop site and the basket load impacted the ground, followed immediately by the 150 foot long line falling straight down onto the basket load. At about the same time, the helicopter entered into a right bank, followed by a left bank. The helicopter subsequently impacted the ground in a steep left bank.

PERSONNEL INFORMATION

The pilot, age 27, held a commercial pilot certificate with airplane single and multiengine land, airplane instrument, and rotorcraft-helicopter ratings. The pilot also held a flight instructor certificate, with airplane single and multi-engine and rotorcraft-helicopter ratings, as well as a mechanic airframe and power plant certificate. On June 24, 2013, the pilot was issued a Class 2 medical certificate, with no restrictions. A review of the pilot's flight records indicated that he had logged over 5,000 hours total flight time, with over 1,800 hours in the make and model of the accident helicopter. During the seven days prior to the accident, the pilot had flown the accident helicopter about 15 flight hours.

AIRCRAFT INFORMATION

The accident helicopter, originally manufactured in 1965 by Bell Helicopter, Inc., was equipped with a Lycoming T53-L-13B engine, composite main rotor blades and BLR Aerodynamic Solutions Tailboom Strakes and Fast Fin System. The Federal Aviation Administration (FAA) certified the helicopter to be flown from the left seat and issued a restricted type certificate (TC) to Tamarack Helicopters, Inc., which authorized agricultural, forest/wildlife, and external load operations.

Following extensive maintenance and refurbishment by Billings Flying Service, the helicopter was released for maintenance test flights on June 5, 2013. At the time of the accident, the helicopter had flown a total of 14,798 hours, which included 41 flight hours after completion of the refurbishment.

The helicopter was operating about 700 pounds (lbs.) below performance limit capability for a 50 foot out of ground (OGE) hover as the pilot approached the survey location. The basic helicopter weight, as documented in the Form B aircraft weighing record, was 5,240 lbs. Total helicopter weight was about 8,000 lbs., based on an estimated 2,000 lbs. basket load, 500 lbs. fuel load, 180 lbs. pilot weight, and 80 lbs. of miscellaneous tools/personal gear. As calculated with the FAA-approved operator manual supplement hover chart, the maximum weight for a 50 foot OGE hover was about 8,700 lbs.

METEOROLOGICAL INFORMATION

The weather observation station at Cortez Municipal Airport (KCEZ), Cortez, Colorado, located about 21 miles to the south of the accident site, reported the following conditions at 0953: wind 300 degrees at 4 knots, visibility 10 miles, overcast clouds at 8,500 feet, temperature 21 degrees Celsius (C), dew point 12 C, altimeter setting 30.25. Density altitude (DA) conditions at the accident site location were about 9,500 feet, based on an accident site elevation of 7,630 feet and a standard temperature lapse rate from the KCEZ observation.

WRECKAGE AND IMPACT INFORMATION

The helicopter impacted into relatively flat, bush covered terrain and was found resting on its left side. The main fuselage had significant crushing to the nose section and left side. The skids exhibited bending to the left forward crosstube consistent with a left bank nose down attitude at ground impact.

Both composite main rotor blades were attached to the main rotor hub through blade bolts and drag braces. Each blade exhibited damage consistent with ground contact while rotating. The main rotor hub was intact except that the pitch horn from one main rotor blade had separated from its main rotor grip consistent with ground impact forces. The mast remained attached to the main rotor hub assembly and to the transmission. The main transmission chip detector was removed and exhibited no debris or chips. No pre-impact anomalies were observed with the main rotor hub, composite blades, or mast.

The flight controls and hydraulics system exhibited extensive damage from impact. The helicopter was equipped with dual flight controls. The pilot was flying from the left seat, with instrumentation on the left pilot door for external load operations. The pilot's cyclic stick and collective stick exhibited fractures consistent with overload near their mounting locations. The rotating controls exhibited continuity to the stabilizer bar and to each pitch change link and pitch horn. The hydraulic pump was removed from the hydraulic drive quill on the sump case and the splined driveshaft was intact. The splined driveshaft was rotated with pliers and minimal resistance occurred.

Engine examination revealed compressor shroud metal spray on the aft side of the 2nd stage power turbine, consistent with operation at the time of impact. Evidence of engine rotation at the time of impact included rotational scoring and tearing on the inner diameter of the particle separator housing assembly, as well as first stage axial compressor blades either broken from the drive hub or exhibiting rotational scoring/leading edge damage.

A follow on airframe examination was conducted at the salvage location. The left and right hydraulic servos were intact and the frame mounts exhibited multiple fractures consistent with overload. Due to the airframe structure damage, many hydraulic lines were fractured due to overload forces. No chafed hydraulic lines were observed.

The hydraulic switch was observed in the ON position. This hydraulic switch, located on the far right side of center pedestal, is out of immediate reach for a pilot flying from the left seat.

Light bulb filaments from the cockpit main annunciator panel, instrument panel annunciators, and pilot's auxiliary annunciator were examined for filament stretch with a USB connected microscope. Both of the bulb filaments for the hydraulic pressure annunciator light, as well as the single bulb filament for the auxiliary master caution light, appeared to be stretched. The remainder of the annunciator system bulb filaments did not appear to be stretched. Filament stretch of the hydraulic pressure annunciator and auxiliary master caution light bulbs was confirmed by the Bell Helicopter engineering laboratory.

Several hydraulic fluid drips were observed on hydraulic lines. A hydraulic check valve near the tail rotor control actuator servo could be rotated by hand between solid hydraulic lines and fittings. The hydraulic check valve and proximate hydraulic pressure line were intact and connected to undistorted walls in the aft belly compartment. In this same compartment, several rags were found soaked with hydraulic oil.

The external cargo hook was tested and responded accordingly when commanded to release both electrically and mechanically. The cargo hook electrical switch and manual cargo hook release could not be tested, due to damage.

MEDICAL AND PATHOLOGICAL INFORMATION

On July 27, 2013, an autopsy was performed on the pilot by a forensics pathologist at the Ertel Mortuary, Cortez, Colorado, as authorized by the Dolores County Coroner. The cause of death was attributed to blunt force injuries. The FAA's Civil Aeromedical Institute in Oklahoma City, Oklahoma, performed toxicology tests on the pilot. No carbon monoxide, ethanol, or drugs were detected during testing.

The pilot was wearing a headset, vice a helmet, during the accident sequence. The left seat, which the pilot was flying from, was equipped with a seat belt and shoulder harness. The seat belt right side attachment fitting was fractured. The shoulder harness was not connected to the seat belt and the straps were not restrained within the guide at the top of the seat.

TESTS AND RESEARCH

Hydraulic system components were examined and tested at the Bell Helicopter engineering laboratory, with oversight by FAA and NTSB personnel. Hydraulic bench testing revealed that the right cyclic, left cyclic, and collective hydraulic servos operated normally and did not exhibit evidence that would preclude normal operation prior to the accident. Testing of the hydraulic control panel, filter, and pump revealed no anomalies or leakage. Of the various hydraulic fixed lines and flexible hoses that were in a condition that allowed for testing, none exhibited evidence of leakages with the exception of two components: a leak was observed with one hydraulic return line near the transmission, as well as a leak to the hydraulic check valve/pressure line to the tail rotor control actuator servo near bulkhead station 211.

The hydraulic return line located near the transmission had 45 degree shear laps in the flareless fitting due to removed tubing material that allowed the sleeve to lap over itself during tightening. It could not be determined if ground impact opened one of the sleeve laps, or if a noticeable 10 degree bend in the tube adjacent to the leak was impact-related.

The hydraulic check valve exhibited longitudinal cracks at both threaded ends. The cracks were a result of stress corrosion cracking, as evidenced by intergranular features, crack branching and corrosion. No material discrepancies were noted. Evidence of Teflon tape was present at the outl...