ERA09LA031
ERA09LA031

HISTORY OF FLIGHT

On October 30, 2008, about 1045 eastern daylight time, a Hughes 369D, N606AS, was substantially damaged during a forced landing near Carroll Township, Pennsylvania. The certificated commercial pilot was seriously injured. Visual meteorological conditions prevailed, and no flight plan was filed for the local rotorcraft external load flight, which was conducted under the provisions of 14 Code of Federal Regulations Part 133.

According to a Federal Aviation Administration (FAA) inspector, the purpose of the flight was to trim trees located adjacent to a power transmission line easement. The helicopter performed this task utilizing an array of saw blades suspended below it.

According to a written statement submitted by the pilot, he departed Washington County Airport (AFJ), Washington, Pennsylvania, about 0800, and proceeded to a landing zone located near the accident site. After landing, the pilot reviewed maps and plans for the operations to be formed with representatives of the power company. After the helicopter was “topped off” with fuel, the pilot proceeded to an area about 8 miles northwest to trim a “danger tree,” then proceeded back to the landing zone. The pilot then cut a line of trees running west to east, on the south side of the power lines, worked back towards the west on the north side of the line.

When the pilot reached a point that was adjacent to the landing zone, he noted a vibration in the helicopter, scanned the instruments, but found nothing of concern. While making a “bottom pass cutting,” the pilot heard a loud “bang” and the helicopter entered a spin to its left. After about 270 degrees of turn, the piloted noticed a large tree, flew to the right of it, then turned to the left, as the helicopter descended rapidly. The pilot lowered the collective pitch control, but did not have time to release the saw suspended below the helicopter or look at the instruments. About 50 feet above the ground, and just above the trees, the pilot increased the collective pitch to slow the descent. The last 30 feet of the descent was a “freefall.”

The pilot additionally stated that prior to hearing the “bang,” he did not notice any caution lights, and did not hear, see, or feel anything suggested an imminent problem.

A witness stated that while driving, she noticed the accident helicopter, and that there appeared to be “fluid coming out of it.” She elaborated that it initially seemed that the helicopter was dumping water on the terrain to extinguish a fire, but the witness did not see any smoke. Shortly after noticing the fluid emanating from the helicopter, the helicopter began to “spin in a counter-clockwise motion” and went straight down and out of her view. The witness estimated that the entire event lasted about 30 seconds.

PERSONNEL INFORMATION

The pilot reported 4,198 total hours of flight experience, with 4,000 hours in the accident helicopter make and model. His most recent FAA third-class medical certificate was issued on January 23, 2008.

AIRCRAFT INFORMATION

The helicopter was equipped with an Allison Engine Company (Rolls-Royce) model 250-C20B engine. Review of maintenance records for the helicopter revealed that at the time of the accident, it had accrued 14,020 total hours of operation, and the engine had accumulated 8,279 total hours of operation. The most recent 100 hour inspection was completed on October 5, 2008, at 13,970 aircraft hours. The engine compressor case-half inspection and the gas generator turbine overhaul were completed 589 hours prior to the accident, and the fuel nozzle flow check was completed 153 hours prior to the accident.

METEOROLOGICAL INFORMATION

The reported weather at Capitol City Airport (CXY), Harrisburg, Pennsylvania, located about 9 nautical miles northeast, at 1056, included winds from 330 degrees at 12 knots, gusting to 18 knots, clear skies, 10 miles visibility, temperature 7 degrees Celsius (C), dewpoint -4 degrees C, and an altimeter setting of 30.43 inches of mercury.

WRECKAGE AND IMPACT INFORMATION

The helicopter was examined at the accident scene by a Federal Aviation Administration (FAA) inspector, and was subsequently transported to a wreckage recovery facility and re-examined on November 13, 2008. The helicopter came to rest upright, and exhibited extensive damage to both airframe and drivetrain components consistent with a vertical impact. The helicopter's engine sustained little visible external damage. The lateral cyclic, longitudinal cyclic, and collective main rotor control linkages exhibited continuity throughout the full range of movement from the cockpit controls to the upper flight controls and the rotor head.

Inspection of the engine revealed evidence of vertical impact. The right side and bottom engine mount struts were fractured. All fuel, lubrication, and pneumatic lines, coupling "B" nuts and their associated fittings were checked by hand for security, and were found at least finger tight and marked with torque paint. The tunnel on the exhaust collector contained an approximate 2-inch tear. Inspection of the inlet plenum chamber did not reveal any evidence of foreign material or missing hardware. Inspection of the compressor inlet showed no signs of erosion, or foreign object damage. Borescope inspection revealed no anomalous findings back to the third stage compressor vanes. Inspection of the fourth stage turbine wheel via the exhaust collector revealed minor foreign object damage. Neither the N1 nor N2 systems could be rotated by hand.

The two aft compressor discharge tube (CDT) snap rings were not engaged. Pre- or Post-accident disengagement of the snap rings could not be determined. Evidence of fretting was found near the base of the CDTs and the air discharge tubes were not fully seated into their respective couplings of the outer combustion liner.

TESTS AND RESEARCH

Turbine Outlet Temperature (TOT) Instrument

The helicopter's aftermarket TOT instrument was removed and forwarded to its manufacturer for functional testing, and download of non-volatile memory. The TOT instrument was subjected to a full acceptance test procedure, and all parameters tested were within prescribed tolerances. Download of the instrument's non-volatile memory revealed that the unit recorded a turbine over-temperature event with a peak temperature of 960 degrees C (1,760 degrees Fahrenheit [F]), and an over temperature condition for a total duration of 20 seconds.

Engine Examination

On February 12, 2009, the engine was examined at the Rolls-Royce facility in Indianapolis, Indiana under the supervision of an FAA inspector. No evidence of foreign object damage or compressor failure were observed. Upon removal from the gearbox, the compressor rotated freely by hand and was not disassembled. The outer combustion case exhibited minor impact damage, but was otherwise unremarkable. No evidence of streaking or other combustion anomalies were observed. The outer combustion liner-CDT mating surface exhibited evidence of an uneven/incomplete seal as well as fretting near the sealing surface. The combustion liner was undamaged and exhibited sooting deposits at the 6 and 12 o'clock positions. An orange room temperature vulcanizing-type (RTV) sealant was found around the sealing surfaces between one of the CDTs and the outer combustion chamber. Some of the sealant was also present just inside the outer combustion chamber. Wear patterns observed on the seal rings of both CDTs were not uniform, and exhibited deep contact rubs in some areas, and no evidence of contact in others.

The turbine section was not free to rotate. The aft side of the first stage nozzle was missing its rope seal. The first stage turbine blades exhibited extensive thermal damage. The second stage had shed a single blade consistent with high cycle fatigue. The third and fourth state turbines exhibited thermal and impact damage, along with a loss of substantial portions of the outer and inner knife air seals consistent with out-of-balance operation following the loss of a turbine blade. The compressor to turbine coupling exhibited deep circumferential rubbing at the turbine end.

The first-, second-, and third-stage turbine wheels, and the nozzles were removed and subjected to detailed metallurgical examination. The tips of the first-stage turbine wheel airfoils exhibited incipient melting and fracture from centripetal forces. Metallographic studies of a first-stage wheel airfoil revealed incipient melting in the ALPAK coating diffusion zone and gamma-prime solutioning of the base metal alloy. Both attributes indicated that the tips of the airfoils were heated to more than 2,100 degrees F. Buckling and cracking occurred along the trailing edges of the airfoils in the first-stage nozzle. Such airfoil deformation was consistent with high, non-uniform operating temperatures. The first-stage blade track exhibited metal spatter deposits with a composition consistent with the first-stage airfoil alloy. Metallographic studies also indicated the occurrence of blade tip rub in the first-stage blade track happened prior to the deposition of the metal spatter from the incipient melting of the first-stage airfoils.

Metrology measurements on the outside diameter of the second-stage nozzle revealed the first and second-stage blade tracks were circumferentially deformed, and had become out of round beyond specification in a manner consistent with thermal exposure. The second-stage nozzle exhibited trailing edge buckling of the airfoils, incipient melting of the airfoils, and reflow of the braze alloy used to join the diaphragm to the nozzle. Metallographic sections of the nozzle in some regions exhibited gamma-prime solutioning of the base metal alloy (which occured above 2,100 degrees F) in the inner and outer bands as well as the full length of the airfoil. However, in other regions of the nozzle, only the center of an airfoil exhibited gamma-prime solutioning of the base metal alloy. This variation in gamma-prime solutioning was consistent with wide a variation in thermal distress of the second-stage nozzle materials and was consistent with non-uniform combustion or burner profile.

One airfoil was fractured by high cycle fatigue on the second-stage turbine wheel. The fatigue crack initiated on the suction side of the airfoil near the mid-chord root. Metallurgical analysis of adjacent second-stage wheel airfoils revealed stress rupture cracks and gamma prime solutioning indicative of excessive temperature exposure. Additionally, metallurgical analysis of an adjacent airfoil tip indicated that tip rub had occurred, and likely preceded the fatigue failure of the fractured airfoil.

ADDITIONAL INFORMATION

A review of Rolls-Royce maintenance manuals revealed no provisions for sealing the combustion liner to CDT interface utilizing an RTV sealant.

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