Eurocopter Deutschland MBB BK 117 N146DU

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 Examination of the wreckage at the accident site revealed that the No. 1 engine twist-grip throttle control in the cockpit was found in the OFF position (which matched the indicator on the No. 1 engine fuel control unit). In order for the control to be placed in this position, the pilot would have had to press a release button on the grip to rotate it below the IDLE position. The No. 2 engine twist-grip throttle control was found in the FLIGHT position (which matched the indicator on the No. 2 engine fuel control unit). This evidence indicated that the pilot likely shut down the No. 1 engine and that the helicopter continued to fly for some time with power being provided only by the No. 2 engine. According to the helicopter’s FLM there were three events that might prompt a pilot to shut down an engine in flight; an engine fire, an engine CHIP indication, or low engine oil pressure. It is possible that the pilot inadvertently shut down the engine No 1. in response to the No. 2 engine ENG CHIP caution. However, the available evidence did not indicate why he might have chosen to shut down an engine rather than reduce the engine power to idle. It is also possible that the pilot erroneously shut down the No. 1 engine after receiving one or more of the above-mentioned cockpit indications, which may have been unexpected and/or confusing. Single Engine Performance Performance calculations indicated that, given the takeoff weight and ambient conditions, the helicopter would have been able to fly to a suitable landing location with only one engine operative without exceeding maximum continuous power. However, the helicopter would not have been able to hover to land, which would have required the pilot to make a running landing onto a smooth, firm surface. The GPS tracking data from the helicopter were not recorded at a sufficient frequency to determine the helicopter's track, speed, and descent profile before impact. As a result, the investigation could not determine if the pilot maneuvered for an immediate landing or a diversion to an alternate location, nor could it be determined if the pilot established the appropriate speed for a one-engine-inoperative (OEI) condition. However, witness reports indicated that the helicopter appeared to be in control with the main and tail rotors turning at an estimated altitude of about 300 ft above the ground with little or no forward speed just before the helicopter’s rapid final descent. Thus, it is possible that the pilot was attempting an emergency OEI landing when the loss of power in the No. 2 engine occurred. While the helicopter’s flight manual provided a procedure for a dual engine failure, it did not provide autorotation performance data to compute height and speed combinations that should result in a successful autorotation with both engines inoperative. Nevertheless, the helicopter might have been at an altitude that was too low and/or an airspeed that was too slow to allow for a successful autorotative landing when the loss of power in the No. 2 engine occurred. Assessment of Potential Fire in No. 2 Engine Compartment Filaments in all four bulbs in the No. 2 engine fire warning light were found stretched, consistent with their illumination at the time of impact. Each engine’s electrically operated airframe fuel shutoff valve was found in the open position. Because each valve would close if its respective fire indication button was pressed, the pilot likely did not press either fire indication button. This suggests that the pilot did not inadvertently shut down the No. 1 engine in response to a potential No. 2 engine fire indication, and that the No. 1 engine shutdown occurred at some point before the illumination of the No. 2 engine fire warning light. With the No. 1 engine already shut down, the pilot would likely not have pressed the No. 2 engine fire indication button, because it would have automatically shut down the helicopter’s remaining source of engine power. Although witnesses reported that during the final moments of the accident that smoke had been trailing behind the helicopter that was black and/or blue in color, the investigation could not determine if the origin of the smoke was from an active engine compartment fire, engine oil exposed to hot engine components, poor combustion in the No.2 engine, and/or another source. Although a plume of smoke can emanate from an engine’s exhaust if the engine is shutdown using the engine fire button, the open airframe fuel shutoff valves suggest that neither fire button was pressed. The investigation could also not determine if any smoke was present in the cockpit and if that could have affected the pilot's visibility inside or outside of the helicopter. Although reports of erroneous fire indications in BK117 C2 helicopters have occurred in the past, based on the fire indication light bulb filament stretching, and the multiple witness reports of smoke emanating from the helicopter in flight, it is possible that an in-flight fire occurred in the No.2 engine compartment; however, the impact damage and post-crash fire precluded a conclusive determination of the presence and origin of an in-flight fire. The postaccident condition of the No.2 engine and the rotor system indicated that prior to impact, the No.2 engine lost all power. The cause of the ultimate complete loss of power in the No. 2 engine could not be determined. Although it is reasonable to consider that it may have been the result of effects of the rear bearing failure, or as a consequence of an unmitigated in-flight fire, there was insufficient evidence to conclude if or how either of these conditions led to the complete loss of power in the No. 2 engine. Possible Indication of Impending Bearing Failure A review of records from engine oil tests performed in the 9 months preceding the accident revealed that metallic contaminants were detected in the No. 2 engine oil at levels considerably higher than in the No. 1 engine. Although the contaminant levels did not exceed those specified by the engine manufacturer to warrant action, the contamination levels fluctuated significantly between tests. In one case, the concentration of iron, which could be an indicator of impending bearing failure, doubled between tests. The oil test evaluation procedures did not include steps to monitor trends of contaminant concentration levels over time. If the engine manufacturer’s procedures had included appropriate trend monitoring criteria, the impending bearing failure in the No. 2 engine might have been detected and mitigated.

Quoted verbatim from the NTSB record.