WPR10GA102
WPR10GA102

HISTORY OF FLIGHT

On January 08, 2010, about 1220 mountain standard time, a Hughes 369D, N500LW, experienced a loss of engine power and landed hard in the wilderness near Kooskia, Idaho. The helicopter was being operated under contract to the Idaho Department of Fish and Game by Quicksilver Air, Inc., as a public flight. The air transport pilot and two passengers sustained serious injuries; the helicopter sustained substantial damage. The local area flight originated from a ranger station about 25 minutes prior to the accident. Visual meteorological conditions prevailed, and no flight plan was filed.

During a telephone conversation with an NTSB investigator, one of the passengers reported that both he and the other passenger were employed by the Idaho Department of Fish and Game. He stated that the purpose of the flight was a wolf capturing mission. While maneuvering around the area where Kelly Creek and Moose Creek intersect, several wolves were spotted. As the pilot maneuvered the helicopter closer to the wolves' location, a loud "bang" emitted from the engine compartment, followed by a loss of engine power and slight vibration. The rugged terrain below was densely populated by trees. With the low rotor rpm light illuminated, the pilot attempted to fly the helicopter toward an opening. The helicopter landed hard on ice.

After being released from the hospital the pilot spoke to an NTSB investigator and recalled that prior to the accident, he was maneuvering the helicopter in a hover about 300 to 400 feet above the ground. He heard a "bang," which was immediately followed by the nose yawing to the left and the engine out light illuminating. He lowered the collective and entered an autorotation.

The pilot further reported that he had a transient chip light illuminate [Engine Chips] for several seconds about one hour prior to the accident. He stated that he performed a check on the enunciator panel at that time and all lights were operational; no other lights illuminated until the loss of power.

HELICOPTER INFORMATION

The Hughes 369D, serial number 470120D, was manufactured in 1977, and had accrued a total time in service of about 13,197 hours at the time of the accident. The owner procured the helicopter in April 2005. The last airframe inspection occurred on December 31, 2009, about 30 flight hours prior to the accident. The helicopter was equipped with a Rolls-Royce 250-C20B Turboshaft engine, serial number CAE-833319, rated at 420 shaft horse power (SHP). The engine had accumulated a total time in service of 7,241.9 hours and had undergone a gearbox repair in October 2009, with the engine being reinstalled on the airframe on November 02, 2009.

The engine's compressor assembly, part number (P/N) 6890550 had accumulated 301.9 hours since overhaul and was installed in the accident helicopter in June 2009, about 6 months prior to the accident. A review of the records revealed that during the overhaul, a new No. 2 bearing was installed: P/N 6889093AL, serial number (S/N) TA 36-0510763; this is the bearing that was believed to be on the helicopter at the time of the accident.

The operator stated that since owning the helicopter there was no welding or high-energy maintenance performed on the helicopter. It was further noted that the helicopter was not used for any hot line powerline work and was never grounded on the ramp except routine fueling standard grounds provided by fueling facilities.

The Rolls-Royce/Allison 250-C20B is a two-shaft turboshaft engine with a combination compressor, which consists of a six-stage axial compressor attached to a one-stage centrifugal compressor. The engine incorporates a reverse-flow annular combustor, a two-stage high-pressure turbine (also referred to as the N1 gas producer turbine), and a two-stage low-pressure turbine (also referred to as the N2 power turbine). The gas path along the Rolls-Royce/Allison 250 engine flows into the inlet, through the compressor's axial and centrifugal stages, into two external air transfer tubes and to the combustor, which is located at the very rear of the engine. The gases then turn 180 degrees toward the front of the engine and proceed through the two-stage gas producer turbine (N1) and the two-stage power turbine (N2). Finally, the gases are directed out of the exhaust collector and upward through two exhaust outlets.

N1, consisting of turbine wheels and nozzles #1 and #2, drives the compressor section of the engine through an inner shaft, while N2, consisting of turbine wheels and nozzles #3 and #4, drives the power output gear (to the main rotor transmission) and the accessory gearbox through an outer shaft. The inner shaft rotates independently within the outer shaft.

TESTS AND RESEARCH

The fuselage sustained significant damage with most damage on the left side. The aft compartment floor was pushed upward approximately 8 inches on the left side. There were multiple main rotor blade strikes to the tail boom, which was separated at fuselage station 240. Examination of the flight control system revealed no evidence of mechanical malfunction or failure. The visual inspection of the fuselage and airframe components during the wreckage inspection did not detect any burn marks, arcing, pitting, or signs of high temperatures stress associated with electrical arcing.

Engine Examination

External examination of the engine revealed that it had sustained crush damage to the combustor case. A speed handle was inserted into the N1 and N2 tachometer drive pads on the gearbox housing. Rotational continuity was established for the N2 gear train and it was found to rotate freely without binding. During the attempted rotation of the N1 gear train, heavy binding was felt. The engine chip detectors were found to contain many metal particles imbedded.

Further disassembly revealed that the No. 2 bearing, P/N 6889093AL, was fractured. The ball bearing retainer was separated in two locations and numerous balls within were gouged on their surface and deformed. A thin oil film was present and no noticeable discoloration was observed. The shim and oil slinger were intact.

The compressor case halves were removed, revealing that the rotor assembly was intact with no evidence of damage observed. The diffuser scroll exhibited circumferential rubbing, consistent with the impeller blades making contact. The impeller blades were slightly rubbed with shiny metal showing and a lip of material displaced opposite the direction of rotation. The No. 1 bearing was intact and had a thin oil film; it was normal in coloration.

The accessory gearbox housing was removed and investigators again attempted to rotate the N1 gear train via the N1 tachometer drive pad noting that the binding was still present. The gears appeared to all be intact with an oil film present. Metallic chips (ferrous) were present throughout the gearbox. Binding was noted during rotation of the pump gear train, but loosened with continuous rotation. Removal of the oil and scavenge pump revealed that the gears were intact with metallic chips present. Light circumferential rubbing was noted on the separator plate (where the gear interfaced).

The turbine section was removed and partially disassembled. The Nos. 6, 7, and 8 bearings were intact and rotated freely in a thin oil film. Molten metal deposits were observed throughout the turbine section. The nozzle shield appeared normal.

The fractured No. 2 bearing was a Federal Aviation Administration (FAA)- Parts Manufacture Approval (PMA) replacement part for Rolls Royce P/N 6889093; it was identified as being manufactured by Timken Alcor Aerospace Technologies, Inc.

Bearing Examination

Examination of the ball retainer at the NTSB Materials Laboratory revealed that it was fractured in several places. The fracture surfaces on the retainer pieces were too damaged to determine the micro mode of fracture. The outer raceway was damaged due to scoring, plastic deformation, and adhesive wear; no as-manufactured surface remained on the raceway.

All 13 balls of the bearing were found, of which many had flat spots consistent with adhesive wear due to sliding contact; some of the balls were reduced in nominal diameter due to wear. The raceway of the inner bearing race halves on the bearing exhibited adhesive wear and deformation damage; no as-manufactured surface remained in the raceway. Due to contact damage, no specific ball track path was observed. Deformation at the edge of the raceway was consistent with excursions of the ball track path to the edges of the raceway.

Optical and digital microscope examination of the side of the No. 2 bearing’s outer race revealed crater-like pitting and the deposition of material. Similar crater and spheroid features were observed on the mating rear diffuser plate bearing housing shoulder. Because the side of the No. 2 bearing outer race and the shoulder on the diffuser bearing housing mate during use, the craters and spheroids appeared as mirror images.

Energy dispersive spectroscopy (EDS) revealed that the craters and spheroids contained deposits of chromium, consistent with material transfer from the diffuser plate bearing shoulder (which had a chromium coating) to the edge of the No. 2 bearing outer race. The spheroids, fine craters, and general as-cast appearance of the crater features were consistent with damage produced by electrical arcing.

Metallic deposits of fused and resolidified metal were found on the surface of the oil slinger (on the oil labyrinth seal side) in a manner typical of splatter. When evaluated by EDS, both the base material and the splatter deposits contained primarily iron and chromium. The morphology of the splatter deposits was consistent with fused metal ejected from the gap between the oil slinger and the oil seal. The ejected metal splatter had a primary direction of orientation, consistent with splatter formation under static rather than rotational conditions.

Caution/Warning System

The caution (yellow) and warning (red) lights consist of separate and removable modular units mounted across the top of the instrument panel. Each light indicator assembly contains four midget, flange base lamps. Caution light circuits are completed to ground through their respective fault sensors. Once the caution light circuit is grounded, the caution light illuminates. The PRESS TO TEST switch illuminates the entire warning and caution light panel when pressed to ensure that the lights are functioning and is part of the engine prestart procedures listed in the Rotorcraft Flight Manual (RFM). All of the modular units are indicators except for the PRESS TO TEST modular unit that functions only as a switch.

The engine chip detector indicator is an amber caution light indicator. The indicator module section of the caution light indicator contains the ledged face and houses four miniature lamps. Access to the miniature lamps is done by pulling the indicator module section away from the switch housing. The switch housing has four spring loaded contacts that contact each of the lamp’s center filaments. The four lamps are grounded through the lamp’s case and are connected to a center ground pin. The ground pin contains a grounding contact collar, spring, and rubber collar.

A postaccident examination of the panel revealed that the "Engine Chips" modular unit did not illuminate when the PRESS TO TEST switch was depressed; however, it illuminated when the actual modular unit was depressed. Disassembly of the modular unit disclosed that the grounding contact collar was pulled back and wedged between the rubber collar, spring, and metallic grounding pin. This condition prevented the indicator module ground from making contact to the switch housing ground unless the indicator module face was pushed and held in. With the indicator module’s grounding collar not in contact with the switch housing ground, the engine chip detector caution light would not illuminate if metallic chips grounded either of the two engine chip detector plugs.

A detailed report with accompanying pictures is contained in the public docket for this accident.

ADDITIONAL INFORMATION

Engine Chips Light Indication

The helicopter's Rotorcraft Flight Manual states that should a pilot encounter an illuminated Engine Chips light, he/she is to land as soon as possible. The applicable chip detectors are then to be removed and inspected for metal accumulation. If chips are found, the engine's Maintenance Manual is to be referenced. The manual does not state a duration the light must be illuminated before checking the chip detector.

In pertinent part, the Maintenance Manual recommends the following action as a result of a magnetic chip light indication:

-Clean the magnetic drain plugs. Perform a 30—minute ground run at power with the rotor turning. Observe engine operating limits and chip warning lights. If operation is normal, remove, inspection, clean, and reinstall all chip detectors. Return engine to service.

Lightning Strike Inspection

According to the helicopter manufacturer, the helicopter's FAA certification basis was under Civil Air Regulations (CAR) 6, which does not include requirements for the protection from lightning strikes. He noted that the 369D model helicopter has had no reported catastrophic damage from lightning strikes and that damage from a lightning strike would result in direct and/or indirect effects to the helicopter that should be detected. Suspect or evidence of a lightning strike on the helicopter requires the Conditional Lightning Strike Inspection.

According to the engine's Maintenance Manual, when a lightning strike occurs, the "aircraft extremities (tail pylon, blades, nose, landing gear, etc.) typically act as points of entry or exit." It specifies that "since the exact electrical path through the aircraft may not be readily traceable following a strike" several actions are recommended.

In pertinent part, it states the following:

-In the event of a lightning strike in the immediate vicinity of the engine(s) (as evidenced by charring, bum marks or pitting associated with electrical arcing on the engine cowl, compartment or inlet) remove the engine(s) prior to further flight.

-Where the aircraft is known or suspected of having been involved in a lightning strike, and entry/exit points either cannot be determined or appear remote to the engine(s), perform the following:

-Inspect the engine compartment for evidence of lightning strike damage.
-Manually rotate N1 and N2 systems and check for binding and abnormal noise. ---Remove, inspect, and clean the engine oil filter.
-Remove, inspect, and clean the magnetic plugs.

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