Incident Overview
On the morning of the occurrence, a Bell 206L helicopter (registration C-FVIX, serial number 45139) was in cruise flight at approximately 700 feet above sea level. The pilot heard a sudden unusual noise and subsequently experienced an engine power loss. He lowered the collective and checked the instruments while scanning for a landing spot. The engine was still running, but the turbine outlet temperature (TOT) rose rapidly and exceeded the gauge range.
Raising the collective slowly caused the main rotor to droop. The pilot advised the two passengers of an engine failure and entered autorotation. During the flared landing, he pulled the collective and confirmed no engine power as the low rotor horn sounded. The helicopter landed on a logging road near Tasu Creek, Queen Charlotte Islands, Sandspit area, at 0829 Pacific daylight time. The pilot shut down the engine immediately. There were no injuries and no airframe damage.
Engine Examination
The helicopter was equipped with a Rolls-Royce Allison 250-C20R engine (serial number CAE 295208). Visual inspection at the site revealed a large crack, about four inches long, in the lower weld joint of the ignitor and fuel nozzle mount at the back of the combustion section. The left-hand exhaust duct was also cracked at a weld on the forward side of the outer combustion case, about six inches down from the top, with impact damage appearing to originate from inside the engine. There was no visual damage from the substantial TOT-indicated over-temperature. A grinding noise was heard when the blades were rotated backwards.
Fuel and oil samples showed no obvious anomalies. The lower chip detector was clean; the upper and freewheel chip detectors contained small amounts of metal. The compressor showed no foreign object damage but could not be turned by hand.
The engine had a total time since new of 6694.4 hours. Technical records indicated an engine discrepancy about 45 hours before the occurrence: a fuel-control unit adjustment due to hot starting. Another entry noted an incorrect engine oil press/temp gauge (C20B model), with replacement parts ordered but no subsequent action recorded. Additionally, the TOT harness and gauge were replaced about 312 airframe hours prior due to erroneous readings.
Teardown Analysis
Following an engine teardown, component parts were examined by the Transportation Safety Board Engineering Branch. The initiating event was the fracture of one blade on the second-stage turbine wheel. The resulting metallic debris caused secondary damage.
The first-stage turbine wheel exhibited many type A and approximately four type B cracks in the blade rim, considered normal in the rim area. Rub damage was observed on several blade tips and on the inner stage seal, covering arcs of about 180°. Fretting wear was noted on the curvic coupling.
The second-stage turbine wheel had one failed blade (accumulated 375.8 hours TTSN out of a 1775-hour service life). Optical examination indicated blade failure due to fatigue cracking with multiple initiation sites in the fillet area on the convex side. Fatigue cracking accounted for about 75% of the fracture; the remainder was an instantaneous overstress rupture. The origin area was almost parallel to the blade axis. The blade following the failed blade was bent opposite to rotation. Significant tip rub was observed on 75% of the blades, with smearing, blued, and mushroomed material. The balance piston seal showed rub damage around approximately 270° of arc, and the inner stage seal had rub damage on about 80% of the arc.
Failure Mechanism
Possible fatigue initiation scenarios included tip rubbing, physical or metallurgical defects, high-cycle fatigue due to vibration (blade flutter), and thermal fatigue cracking. Tip rubbing was ruled out as an initiating event due to the absence of significant rubbing on the adjacent blade tip and acceptable dimensional analysis of the second-stage nozzle. Metallurgical analysis revealed no manufacturing or material anomalies. Fatigue striation spacing and initiation site indicated the initial mode was low-cycle fatigue cracking, which then progressed in high-cycle mode once cracking started.
Thermally induced fatigue cracking occurs when rapid expansion in the turbine wheel rim produces compressive hoop stresses, leading to localized yielding, followed by tensile stresses at steady-state temperatures that initiate cracking. The location and orientation of cracks in the second-stage turbine wheel were considered the result of thermal fatigue.
Instrumentation and Hot Starts
The engine oil pressure and temperature gauge had incorrect markings (C20B model). The TOT harness and gauge were replaced about 312 airframe hours before the occurrence due to errors. It could not be confirmed whether the ultimate temperature (927°C) for a turbine inspection was reached during hot starts, or if the maximum continuous temperature (810°C) was exceeded. The TOT harness may have had a 0.03 ohm discrepancy, and the incorrect gauge may have indicated 15°C low for 63.1 hours. The effects of these errors on turbine wheel cracking could not be determined, but their contribution was considered unlikely because gamma prime phase distribution was uniform, though short-term overheating could not be ruled out. Hot starting events are not recorded by the helicopter's instrumentation and may not be accurately recorded by the operator. About 45 engine hours before the occurrence, the fuel control unit was adjusted due to hot starting, but the temperature exceeded was not recorded.
Other Finding
Approximately 25% of the major diameter seal was missing from the rear support due to a bond failure, likely causing a slight loss of engine efficiency.