N449MCRobinson Helicopter R44 II · 2012

On November 13, 2015, at 1648 central standard time, a Robinson Helicopter model R44II, N449MC, was substantially damaged during an in-flight collision with trees and terrain near Alpine, Texas. The pilot and one crewmember sustained serious injuries; two additional crewmembers sustained minor injuries. The helicopter was registered to and operated by Concho Aviation LLC under the provisions of 14 Code of Federal Regulations Part 91 as an aerial survey flight. Day visual meteorological conditions prevailed for the flight, which was not operated on a flight plan. The local flight originated from a private ranch located about 17 miles south of Alpine, Texas.

GPS data suggested that two flights were conducted before the accident flight. The first flight departed about 0830 and concluded about 1051. The data indicated that the helicopter tracked eastbound momentarily before proceeding toward the west and departing the takeoff/landing area. The second flight departed about 1106 and concluded about 1320. The data indicated that the helicopter tracked eastbound and continued in that direction as it departed the takeoff/landing area.

Crewmembers reported that during one takeoff, the helicopter was unable to climb high enough to clear the power lines that bordered the takeoff/landing area on three sides. The pilot subsequently turned around and departed to the west; the direction in which there was no power lines to clear. However, there was a one-story building in that direction. One crewmember estimated that the landing skids cleared the building by about 15 feet.

In addition, according to one crewmember, the pilot had commented that during the takeoff on the morning flight, the helicopter "wouldn't go." The pilot explained that the helicopter had full fuel and four occupants; the low rotor speed warning came on, which can occur if the rotor blades reach full pitch for the given conditions. The pilot added that when that situation occurs the rotor speed needs to be increased, which can be accomplished by increasing the helicopter's airspeed or initiating a descent.

The crewmembers noted that they were conducting wildlife surveys and predator control at the time of the accident. They stated that the pilot was maneuvering when the low rotor speed warning activated. The pilot reportedly attempted to regain control by descending into a small canyon; however, his efforts were unsuccessful and the helicopter subsequently impacted some trees and then the ground. The pilot did not have any recollection of the accident sequence itself.

GPS data revealed that the accident flight departed at 1515; the final data point was recorded at 1648. During takeoff, the helicopter initially tracked eastbound, before reversing course toward the west momentarily, and then ultimately departing the area to the east. The helicopter proceeded north and west of the departure point with multiple turns and course reversals. The maximum extent of the flight was 3.00 miles north-northwest of the departure point; the accident site was located 2.31 miles north-northwest of the departure point. GPS altitudes ranged from 4,504 feet to 5,357 feet. The elevation of the departure point and the accident site were approximately 4,507 feet and 4,763 feet, respectively.

Shortly before the accident, the helicopter completed a 360-degree turn at about 4,800 feet and at a speed of approximately 20 knots. Immediately after completing the turn, about 20 seconds before the end of the data, the helicopter entered a gradual left turn from an altitude of 4,805 feet to 4,763 feet. The accident site terrain consisted of rolling hills, low brush, and vegetation.

The pilot held a commercial pilot certificate with helicopter, instrument helicopter, and single-engine land airplane ratings. His airplane rating was limited to private pilot privileges. The pilot also held a flight instructor certificate with a helicopter rating. He was issued a second class airman medical certificate on October 15, 2014, with no limitations or restrictions. The pilot's logbook included an endorsement to act as pilot-in-command of a Robinson R44 helicopter, dated February 14, 2014, in accordance with the requirements of FAA Special Federal Aviation Regulation (SFAR) No. 73 – Robinson R-22/R-44 Special Training and Experience Requirements.

According to the accident report submitted by the operator, the pilot had accumulated 735 hours total flight time, with 697 hours in helicopters and 228 hours in Robinson R44 II helicopters. Of that flight time, about 90 hours was flown within the preceding 30 days and about 12 hours within the preceding 24 hours. All of that flight time was in Robinson R44 II helicopters. The pilot completed a flight review on September 27, 2014, in a Robinson R22 Beta II helicopter.

SFAR No. 73 required that a pilot complete a flight review in a Robinson R44 helicopter within the preceding 12 calendar months in order to act as pilot-in-command of such an aircraft. In addition to meeting the requirements of 14 CFR 61.56, this flight review must include the specific abnormal and emergency procedures flight training specified in the SFAR.

The accident aircraft was a Robinson R44 II helicopter, serial number 13340. It was a four-place design, with a two blade, teetering main rotor system and fixed skid-type landing gear configuration. Anti-torque and directional control was provided by a two blade, teetering tail rotor system. Conventional cyclic/collective controls were utilized and the main rotor controls were hydraulically boosted in order to minimize control feedback forces.

A Lycoming IO-540-AE1A5 engine, serial number L-34923-48E, powered the helicopter. The engine was a six-cylinder, horizontally opposed, normally aspirated, air cooled, fuel injected design. It was capable of producing 260 horsepower at 2,800 rpm. However, the engine installation was derated to 245 horsepower for takeoff (5 minutes) and 205 horsepower for continuous operation. Engine cooling was provided by a direct-drive blower ducted through a fan shroud and engine baffling. Engine power is transferred to the rotor drive system through four V-belts. An electric actuator adjusts the position of the upper sheave to maintain proper belt tension during flight.

According to the maintenance records, the most recent annual inspection was completed on April 8, 2015, at 502.91 hours total airframe time. R44 Service Bulletin SB-89 was complied with on April 26, 2015. An annual/100-hour engine inspection was completed on November 6, 2015, at 605.9 hours total airframe/engine time.

Weather conditions recorded by the Alpine-Casparis Municipal Airport (E38), located about 16 miles north of the accident site, at 1635, were: wind from 160 degrees at 10 knots, 10 miles visibility, clear sky, temperature 20 degrees Celsius, dew point 1 degree Celsius, and altimeter 30.22 inches of mercury. The published elevation at E38 was 4,514 feet. The corresponding density altitude was about 5,770 feet.

The accident site consisted of sparsely-wooded, rolling hills, with low vegetation and brush. The helicopter came to rest on its left side. The main rotor blades remained attached to the hub. The tail boom was separated, fragmented and located at the accident site. The tail rotor blades remained attached to the hub. The landing skids remained attached to the fuselage; however, they were deformed and fragmented.

The NTSB did not travel to the accident site. A postaccident examination was conducted after recovery of the helicopter.

The cockpit/cabin area was compromised, with more extensive damage on the left side of the aircraft. The center and aft fuselage structure appeared intact; however, the fuselage skin was deformed and buckled. The forward tail boom structure remained attached to the upper frame; however, the tail boom structure was deformed and separated immediately aft of the mounting point. The tail boom was separated into six sections, including the forward section that remained attached to the aft fuselage. The tail boom sections were deformed and exhibited scuff marks consistent with main rotor blade strikes. The horizontal and vertical stabilizers remained attached to the aft tail boom section. The upper vertical and the horizontal stabilizer exhibited buckling and deformation damage. The top of the upper vertical stabilizer also exhibited crushing damage.

The main rotor mast and fairing appeared intact. The engine-drive train V-belts appeared intact and properly installed on the engine and drive train sheave assemblies. The main rotor gearbox, main rotor mast, and hub appeared intact. Rotation of the drive shaft produced a corresponding rotation of the main rotor mast, confirming gearbox continuity. The sprag clutch/free wheel unit rotated smoothly and functioned properly. The main rotor blades exhibited gradual bending over the span of the blade, with localized deformation and creasing.

The cockpit cyclic control remained attached to the dislocated floor structure. Forward-aft cyclic control continuity was confirmed from the cyclic stick to the main rotor mast. The left-right cyclic torque tube was separated consistent with overstress failure. Left-right cyclic control continuity was confirmed from aft of the separation to the main rotor mast. The collective control appeared intact, with the exception of an overstress separation at the aft end of the tube. Separations in the main rotor mast fork assembly and main rotor blade pitch change links appeared consistent with overload failures. The tail rotor/anti-torque pedals were deformed, but remained attached. The push-pull tube assemblies associated with the tail rotor/anti-torque control system in the cockpit area was separated consistent with overstress failures. Control continuity was continuous from aft of the cockpit area separations to the aft fuselage. The tail rotor/anti-torque control push-pull tube assembly common to the aft fuselage and tail boom was separated into six segments. The forward and aft segments remained attached to the corresponding bell crank assemblies. The tube segments were deformed and each of the fracture surfaces appeared to be consistent with overload failures.

The tail rotor gearbox remained attached to the tail boom and appeared intact. Rotation of the tail rotor shaft producing corresponding rotation of the input drive shaft, with no binding noted. The tail rotor blades remained attached, and the pitch change links and bellcrank were intact. The tail rotor drive shaft was separated at the forward and aft flex-couplings. Both flex-couplings were deformed and the fracture surfaces appeared consistent with overstress failures. The drive shaft was separated into two sections. The fracture surface associated with the aft section was deformed and appeared consistent with an overstress failure. The forward section of the tail rotor drive shaft was not with the wreckage during the post-accident examination and was presumably not recovered.

The engine remained attached to the airframe engine mount and appeared to be undamaged. The cooling baffling was intact. The engine and drive train were rotated via the fanwheel, with the V-belts in place. The lower spark plugs were removed; they exhibited normal combustion signatures. Internal engine, valve train, and accessory section continuity was confirmed via crankshaft rotation. Compression and suction were obtained at all cylinders. The magnetos remained secured to the engine. Magneto timing was within specifications. The magnetos were removed; both magnetos produced a spark at all leads when rotated. Fuel flow was confirmed from the fuel pump output line during engine rotation. The fuel injection servo remained attached to the engine and appeared undamaged. The engine throttle and mixture control cables remained attached to the fuel injection servo. Bench testing of the servo did not reveal any anomalies consistent with a loss of engine power. Fuel flow values measured within specification limits at high power settings. Disassembly of the fuel servo revealed a small amount of non-ferrous debris on both the metered and un-metered sides of the diaphragm. The fuel strainer and vent screen were intact and free of debris or contamination.

Information obtained from the pilot, crewmembers and operator was used to compile a weight and balance calculation of the accident flight. Based on the information available, the gross weight of the helicopter at takeoff was approximately 2,465 lbs., with longitudinal and lateral center-of-gravity locations of 94.5 inches and 0.03 inch, respectively. The maximum gross weight was 2,500 lbs. The center-of-gravity locations were within the loading limits specified in the pilot's operating handbook.

Performance data related to the out-of-ground effect (OGE) hover ceiling indicated that under the current loading and atmospheric conditions, the OGE hover ceiling was approximately 4,100 feet pressure altitude – about 4,400 feet mean sea level under the current conditions.

The accident flight was being operated by Concho Aviation, LLC; a helicopter services company based in Sterling City, Texas. Because the accident flight was being operated for the purposes of aerial survey, specifically for wildlife survey and predator control, the operator was not required to comply with the requirements of 14 CFR Part 119 or 14 CFR Part 135. The helicopter occupants, in addition to the pilot, were classified as crewmembers rather than passengers for regulatory purposes. The basic requirements of 14 CFR Part 91, General Operating and Flight Rules, governed the flight.

The Pilot's Operating Handbook, Section 3 – Emergency Procedures, provided guidance related to a low rotor speed condition. Specifically, the POH stated:

LOW RPM HORN & CAUTION LIGHT – A horn and an illuminated caution light indicate that rotor RPM may be below safe limits. To restore RPM, immediately roll the throttle on, lower collective and, in forward flight, apply aft cyclic. The horn and caution light are disabled when the collective is full down.

As noted in the FAA Helicopter Flying Handbook, hover performance varies depending on whether a helicopter is operating in or out of ground effect. Ground effect increases the efficiency of the rotor system and reduces the power required to hover in ground effect. Rotor efficiency is increased by ground effect to a height of about one rotor diameter (measured from the ground to the rotor disk) for most helicopters. Above this height the beneficial effects of ground effect are not available; more power is required to hover out of ground effect as compared to in ground effect.

Effective translational lift (ETL) is achieved between 16 and 24 knots horizontal movement of the helicopter or surface wind. Once above the ETL speed, the rotor system completely outruns the recirculation of old vortices and begins to work in relatively undisturbed air. As a result, the rotor system operates more efficiently.