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N5105N accident description

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Crash location Unknown
Nearest city Puako, HI
19.975000°N, 155.843889°W
Tail number N5105N
Accident date 10 May 1997
Aircraft type Hughes 369D
Additional details: None

NTSB Factual Report

On May 10, 1997, at 1620 hours Hawaii standard time, a Hughes 369D, N5105N, en route to Waimea, Hawaii, experienced an in-flight loss of power, collided with hilly terrain, and burned near Puako, Hawaii. The aircraft was destroyed; and the pilot and one passenger sustained fatal injuries. The remaining two passengers received serious injuries. The aircraft was being operated as a personal flight by the pilot/operator when the accident occurred. The flight originated near Mano Point, Hawaii, at 1605. Visual meteorological conditions prevailed at the time and no flight plan was filed.

A witness at the accident site reported that the aircraft was approaching him. He estimated the aircraft was on a northwesterly heading, about 100 to 200 feet agl, when he heard the engine quit. Immediately after the power loss, the aircraft banked left and descended, nose low, until the main rotor blades struck the ground. The aircraft continued forward until the lower forward fuselage also contacted the ground. The aircraft then yawed right about 90 degrees, rolled onto its left side, and then rebounded back into the air. It continued another 20 feet before coming to rest on its left side. The tailboom was severed about 3 feet forward of the tail rotor gearbox and came to rest ahead of the main wreckage. All other debris was found within a 30-foot radius of the main wreckage. The long axis of ground scar from the initial fuselage contact was on a 010- degree magnetic bearing. Winds at the time were reported to be from 130 degrees at 8 knots.

The center front seat and right front seat passengers were able to exit the aircraft after the crash. The pilot was attempting to assist the remaining passenger, who was sitting in the right rear seat, when the fire spread to the cabin.

A postcrash fire erupted about 30 seconds after the aircraft came to rest. Two auxiliary fuel cells are located below the waterline in the forward fuselage. Evidence of a fuel spill was later found in the vicinity of the point at which the lower forward fuselage first struck the ground. After the fire erupted, the aircraft was consumed by fire from fuselage station (FS) 124.0 forward and below the main rotor system.

The cockpit instruments and warning systems were destroyed in the fire. The restraint system lap belts and shoulder harness straps were also consumed by the fire. The buckles were located and the belt attachment fittings were found still connected to their anchor points.

The engine sustained both heat and impact damage. The outer combustion case was buckled, the engine mounts were buckled and fractured, and the throttle-to-fuel and collective-to-governor control rods were bent.

The fuel nozzle was removed and inspected by Federal Aviation Administration (FAA) airworthiness inspectors with no anomalies noted. The upper and lower magnetic chip detectors were found to be clean. The No. 1 and No. 2 sections did not rotate, and there was 360-degree outward metal displacement on the compressor case halves in the vicinity of the third stage wheel.

Disassembly of the compressor revealed extensive damage to the compressor wheels and stator vanes. Metal particles were found in the compressor case. All the blades from the third, fourth, and fifth stage compressor wheels were sheared from the rotor. The sixth stage blades were partially sheared. The first stage blades showed heat damage and discoloration, while the second stage showed trailing edge damage. Damage to the compressor case corresponded to the damaged compressor wheels, except for the sixth stage stators that were sheared completely.

The combustion section was disassembled and metal particles were found. The metal was in the form of loose particles, as well as having been deposited on other internal parts as molten metal.

The compressor halves, compressor rotor, and metal particles were forwarded to the Safety Board Materials Laboratory for metallurgical analysis. The complete laboratory report is appended.

A computation of the aircraft weight and balance by FAA inspectors and a representative of the manufacturer revealed that the aircraft center-of-gravity was 98.02 inches, exceeding the forward longitudinal limit of 99.0 inches. The manufacturer's representative stated that this condition limited the amount of aft cyclic control available and limited its effectiveness in an autorotative deceleration.

The aircraft logbooks are believed to have been destroyed in the fire. It was estimated that the aircraft had flown about 60 hours since its annual inspection. This estimate was based on conversations with maintenance personnel who were familiar with the operation of the aircraft.

The engine was overhauled in 1991 with a total engine time of 1,461 hours. It had operated an additional 3,296 hours since overhaul at the time of the last annual inspection.

During the annual inspection, the engine had been removed to comply with Allison Customer Service Letter (CSL) 1172, Compressor Case Erosion Inspection. In the course of the inspection, foreign object damage (FOD) was found on the first, second, third, and sixth stage compressor wheels. The compressor case halves were split and the damaged blades were blended in accordance with Allison maintenance manuals. The first stage turbine was visually inspected and the fuel nozzle was replaced.

An inspection by a representative of the Office of Aircraft Services (OAS), Department of the Interior, was made as part of the application process for a government contract. The inspector found the No. 1 gauge to be inoperative. The inspector also identified the lack of historical information indicating the total times on the governor, fuel pump, fuel nozzle, bleed valve, and fuel control. Finally, he noted an unusual rubbing sound during coast down. Investigators were unable to find any records indicating that the No. 1 gauge had been repaired or replaced.

The Safety Board metallurgist found that one of the second stage vanes leading edge contained erosion damage. The leading edge erosion damage was part of a larger band of erosion that extended between the leading and trailing edges on the pressure side of the blade. The fracture surface contained mechanical damage that extended between the eroded leading edge and an area located approximately 0.07 inch aft of the eroded leading edge. Fatigue cracking, with propagation in the aft direction, was found on the fracture surface that extended from this damage to an area located approximately 0.28 inch aft of the eroded leading edge.

The thickness of the second stage vane in the eroded area was approximately 0.26 inch. The Allison 250-C20 Series Operation and Maintenance manual states that vanes eroded below a thickness of 0.28 inch at midchord cannot be returned to service.

A visual inspection of the revealed that the first and second stage blades were intact, although the trailing edge of the second stage blades exhibited deformation damage. Blades from the stages three through five separated near the root while the sixth stage blades separated approximately midway point of the blade span. The fracture of the blades from the fourth through sixth stages exhibited rotational damage that obliterated the original fracture surfaces. Most of the fractures of the blades from the third stage, however, showed relatively undamaged fracture surfaces.

A microscopic examination of the third stage blade fractures revealed that a portion of these blades contained fatigue cracking that was initiated from the convex side of the blades. Fatigue propagation was away from the convex surface and extended 30 percent through the thickness of the wall. Analysis of the metal particles found in the engine exhibited the same spectra of elemental components as the blades.

Imua Air Service was a sole proprietorship with no employees.

NTSB Probable Cause

erosion of second stage stator vanes, inadequate maintenance inspection for the erosion, and subsequent fatigue failure of a stator vane, which resulted in loss of engine power and a forced landing on mountainous/hilly terrain. Also causal was: the pilot's loading of the aircraft in such a manner that exceeded the forward CG limit, which resulted in his failure (or inability) to properly flare the helicopter during a forced autorotation and landing. The rising terrain and tailwind condition for landing were related factors.

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