ERA17MA316

A failure of the rear bearing in the No. 2 engine, which (1) created multiple and likely unexpected and confusing cockpit indications, resulting in the pilot's improper diagnosis and subsequent erroneous shutdown of the No. 1 engine, and (2) the resulting degraded the performance of the No. 2 engine, until it ultimately lost power. The complete loss of engine power likely occurred at an altitude and/or airspeed that was too low for the pilot to execute a successful emergency autorotative landing.

The pilot was conducting an air ambulance flight to transport a patient to another hospital located about 130 nautical miles away. About 8 minutes after takeoff, at a GPS altitude of about 2,500 ft mean sea level (msl) and a groundspeed of about 120 knots, the helicopter began a left turn toward the south. Although the precise timing and order of events could not be determined, the No. 2 engine experienced a bearing seizure; the engine continued to run. This likely resulted in several cockpit indications described below. It is likely that the pilot then errantly shutdown the No. 1 engine and continued to fly for a brief period utilizing the No. 2 engine. About 1 minute after the left turn began, the last data point was recorded, which indicated that the helicopter was at a GPS altitude of about 1,200 ft msl and a groundspeed of 75 knots. It is likely that the No. 2 engine subsequently lost all power. The helicopter then impacted a shallow turf drainage pathway between fields of tall grass on a farm, and a postcrash fire ensued, which consumed most of the helicopter structure. The lack of any ground scars leading toward or away from the main wreckage indicated that the helicopter was in a nearvertical descent before impacting the ground. One rotor blade was found intact resting in undisturbed 8-ft-tall grass, consistent with little or no rotation of the main rotor system. Neither engine exhibited damage consistent with rotation at the time of impact. Detailed examination of the No. 2 engine revealed that its gas generator shaft rear bearing was seized and damaged. Specifically, all the roller elements were flattened and none of the roller elements would rotate. Several bearing components showed damage consistent with friction between the seized rollers and the inner race and ensuing overheating. These signatures were not observed in the No. 1 engine gas generator shaft rear bearing. The lack of rotation of the roller pins and the damage to the gas generator spool indicated that the No. 2 engine’s rear bearing had failed during the accident flight. The No. 2 engine’s oil return strainer/chip detector was partially obstructed with crystalline carbon-like and metallic debris, and debris was found downstream of the strainer in the oil pump inlet, consistent with some oil flow through the normal path exiting the rear bearing housing. The No. 2 engine 3-way deck fitting was unobstructed; however, it had been exposed to significant heat that could have decomposed the elastomeric tubing and any excess shrink tubing material, if it had been present and blocking the fitting’s ports. Excess shrink tubing material was found obstructing the ports in the No. 1 engine 3-way deck fitting. No liquid oil remnants were found on engine No. 2 components to indicate that engine oil had migrated through the engine’s main air path and through the exhaust. Had oil been present, it might have been consumed by the postimpact fire. The root cause of the No. 2 engine’s bearing failure could not be determined due to the damage it sustained while continuing to operate before impact, and due to impact and subsequent post-crash fire damage. Accident Sequence and Cockpit Indications A lack of recorded flight data precluded determining the precise timing, duration, and order of each event that likely occurred during the accident flight, to include the cockpit indications provided to the pilot; however, based on available evidence, it is possible that the pilot may have encountered the following indications during the final minutes of the flight. Engine Chip Detector Indication Metallic debris found in the No. 2 engine oil return strainer/chip detector downstream of the rear bearing housing likely resulted in an ENG CHIP message on the helicopter’s Caution and Advisory Display (CAD). The helicopter’s flight manual (FLM) listed two procedures for engine chip detection: either shut down the affected engine or slowly reduce power to idle on the affected engine and monitor the engine parameters. The second option was intended to allow the pilot to potentially use the affected engine during landing. Engine Parameter Discrepancy Indication A cockpit display simulation prepared by the airframe manufacturer and an analysis prepared by the engine manufacturer revealed that, during the degradation of the rear bearing, the No. 2 engine experienced an increase in turbine outlet temperature (TOT). This likely triggered an engine parameter discrepancy (ENG PA DIS) caution message that would have appeared in both the left (engine 1) and right (engine 2) columns on the CAD, indicating that a discrepancy was detected in one of the parameters between engine Nos. 1 and 2. The CAD message would not have indicated which parameter had a discrepancy; however, the affected parameter’s numeric value (TOT in this case) would change from white to yellow on the First Limit Indicator (FLI), described below. The FLM procedure for the ENG PA DIS message was, “do not try and match needles, avoid using maximum power, compare the numeric values on the FLI to verify the affected parameter, and land as soon as practicable.” The FLM procedures did not request the pilot to shut down the engine. Other Indications Additional caution messages may have also appeared on the CAD. If a difference in torque between the two engines was detected at greater than 15%, a VAR NR caution message would appear in the center “MISC” column of the CAD, advising the pilot to manually match the engines’ torque values. If a difference in N1 between the two engines of greater that 10% was detected, the ENG SPLIT caution message would appear in both engine columns of the CAD. The FLM procedure for an ENG SPLIT caution message was to adjust the collective lever to one-engine inoperative (OEI) limits or below, turn off bleed air consumers, and analyze engine conditions. First Limit Indicator Display During Bearing Failure A simulation of the primary engine display instrument, the first limit indicator (FLI), revealed that as the bearing degraded, the FLI might have presented data in a way that was unfamiliar to the pilot, possibly causing confusion. Specifically, as the bearing failed, the FLI needle for the No. 2 engine would have changed from indicating torque (TRQ), to indicating turbine outlet temperature (TOT) due to a sudden rise in TOT in the No. 2 engine. Such a switch would have been unusual, because the needles normally reflect TOT during engine start only. The change in the position of the No. 2 FLI needle would have resulted in a large split between both needles. In normal cruise flight, with the FLI needles both representing TRQ, a large split during cruise flight would indicate a difference in TRQ between the engines; thus, the pilot may have erroneously thought that the split was showing that the No. 1 engine was producing much less TRQ than the No. 2 engine, which might have contributed to his decision to shut down the No. 1 engine. The FLI should have indicated, in the numeric section of the display, that the No. 2 needle was indicating TOT, and if appropriate, that the No. 2 engine’s TOT had reached its limit. Despite the split needles, the numeric values for each engine’s TRQ may have at least initially been similar, which could be confirmed by cross-checking the triple tachometer located above the FLI on the instrument panel. The specific condition of the FLI needles showing one engine limited by TRQ and the other engine limited by TOT during cruise flight was not reviewed or practiced as part of the operator’s or the helicopter manufacturer’s training programs. However, depending on operating conditions, the engines could be limited by TRQ, N1, or TOT, which was covered in those training programs, as were engine failures and typical “needle split” conditions that occur during an engine failure. Shutdown of No. 1 Engine

The pilot had been employed with AMC since August 2009. He was the lead pilot and the safety officer at AMC’s JNX base and an AMC maintenance test pilot in the BK117 C2 helicopter. He was also current and qualified on the twin-engine Airbus EC135 helicopter, in which he had accrued 1,100 hours of total flight experience. Before his employment with AMC, the pilot flew twin-engine Sikorsky UH-60 helicopters for the US Army, accruing about 2,300 flight hours, and the EC135 helicopter for another helicopter air ambulance provider. AMC training records indicated that the pilot had completed all required training with no deficiencies. During the pilot’s most recent recurrent training and checkride for the BK117 C2, the pilot performed one-engine-inoperative (OEI) flight procedures and a simulated OEI landing. Recurrent training typically included autorotations, which were practiced to a power recovery at a 3-ft hover, but no autorotations were specifically documented in the pilot’s recurrent training records. At the time of the accident, AMC did not have a BK117 C2 simulator training program, which would allow for practice autorotations to touchdown. According to his training records, the pilot was familiar and current with the indications associated with autorotation and OEI conditions as well as for the behavior of the aircraft. (AMC had been developing a BK117 C2 simulator training program at the time of the accident, which was subsequently implemented). Simulated OEI landings were performed in the aircraft by utilizing power limits representative of OEI performance. Engine fire light procedures were discussed during the training and were the subject of oral questions during the checkride. The pilot’s most recent EC135 simulator training included OEI recoveries, OEI landings, and engine fire light procedures. The indications and procedures in the EC135 are similar to what is seen and performed in the EC145. The pilot's work schedule included 12-hour workdays from 0800 to 2000, with a 6-days-on/6-days-off format. The accident occurred on the third flight leg of the second day of the pilot’s work schedule. According to his wife, the pilot had no issues with his sleep during the 3 days preceding the accident. He was in good health and was not taking any medications. She further reported that he was happy with his life and did not have any major life stressors. The pilot was not employed outside of AMC, enjoyed his job with the company, and had not mentioned any concerns about the company or its helicopters. The pilot's coworkers and managers provided positive feedback about his performance. He was described as professional, well prepared, thorough, and team oriented, and he exhibited good pilot skills. The helicopter impacted a shallow turf drainage pathway, which was about 30 ft wide and 2,000 ft long, located between two fields of 8-ft-tall grass near a wind turbine farm. The fuselage came to rest in a 7-ft-wide ditch in the center of the pathway and was oriented on a magnetic heading of 261°. There were no ground scars leading to or from the main wreckage. Examination of the wreckage revealed that all of the major helicopter components were present at the accident site. The cabin had collapsed downward and was partially consumed by a postcrash fire. The tailboom remained largely intact. Flight control continuity was established from the cockpit area to the rotor systems and engines. The four main rotor blades and the two tail rotor blades remained attached to their rotor hubs. The No. 4 main rotor blade was found rotated about 180° in its hub with the pitch links fractured and partially melted. The outboard 4 ft of the No. 3 main rotor blade came to rest in the 8-ft-tall grass adjacent to the drainage path, and the grass on both sides of the blade was undisturbed. None of the main or tail rotor blades exhibited leading edge damage, chordwise scratches, or other evidence of rotation. The tail rotor shaft remained attached to the transmission, which could not be manually rotated. The portion of the warning unit in the cockpit that contained the No. 2 engine fire warning light/button was located in the wreckage. Examination by the National Transportation Safety Board’s (NTSB) Materials Laboratory in Washington, DC, revealed that the filaments in all four of the No. 2 engine fire warning light/button’s light bulbs were stretched. The portion of the warning panel containing the No. 1 engine fire warning light/button was identified at the accident scene but was subsequently separated from the remaining section of the warning panel during recovery. The No. 1 engine fire warning light bulbs were not examined. No foreign object damage was found on the axial compressor blades of both engines. No damage was observed on the visible portions of the turbine blades at the aft part of the engines. The gas generator of the No. 1 engine moved freely when manually rotated, whereas the No. 2 engine gas generator did not rotate. The helicopter was also equipped with engine throttle twist grips on the pilot's collective control stick. Each twist grip had a lockout button that prevented the grip from being inadvertently rotated from FLIGHT to IDLE and from IDLE to OFF; the button had to be pressed to rotate the grip. The No. 1 engine grip was located at the top of the collective control tube, and the No. 2 engine grip was located immediately below the No. 1 grip (and closer to the hinge of the collective tube). Each grip had a different grooved pattern to manually distinguish one from the other. The No. 1 engine twist-grip throttle control was found in the OFF position. The No. 2 engine twist-grip throttle control was found in the FLIGHT position. The No. 1 engine fuel control unit was found in the 0° (cutoff) position. The No. 2 engine fuel control unit was found in the 62° position, which was slightly beyond the 52° (flight) position. The No. 1 engine fuel shutoff valve was found in the open position. The No. 2 engine fuel shutoff valve was damaged, and its position could not initially be determined. X-ray images of the valve by the NTSB’s Materials Laboratory revealed that the valve was in the open position. Engine No.1 disassembly and component examination did not reveal any discrepancies other than damage due to the crash and the post-crash fire. The No. 1 engine rear bearing oil return strainer/chip detector was absent of debris. The No. 2 engine disassembly revealed that the gas generator shaft rear bearing was mechanically damaged. A detailed examination of the bearing at the engine manufacturer’s laboratory revealed that all of the bearing roller elements were found seized (that is, none of the roller elements would rotate), with the outer bearing race, and had rubbed against the inner rotating bearing race. The roller elements appeared ground down (flattened) and overheated. All of the oil supply pipes and restrictors and jet were found clear. The tubes to and from the rear bearing chamber (the oil supply tube, scavenge tube and vent tube) each contained a thin layer of coked oil, and were not obstructed. Figure 2 compares the No. 2 engine rear bearing with the undamaged No. 1 engine rear bearing. Turbine components and the end of the No. 2 engine gas generator shaft exhibited rotational non-uniform damage. This damage was consistent with some continued rotation of the gas generator spool after the bearing had seized. (The turbine shaft supported by the rear bearing rotates at speeds up to 53,500 rpm.) Figure 2 - Gas generator shaft rear bearings. The No. 2 engine rear bearing oil return strainer/chip detector had carbon-like and ferrous debris in the strainer. Some debris particles were found bridging the gap between chip detector electrodes. The strainer was not completely obstructed by the debris. Downstream of the rear bearing’s casing, the oil return pipe and the suction stage