Viking Aviation Bell 407 incident

The helicopter exhibited multiple impact marks and penetrations in the exhaust duct, exhaust collector, and horizontal fire shield, all consistent with turbine debris exiting out the back of the engine through the exhaust as well as radially outward at the gas producer support case. The gas producer stage 1 turbine wheel experienced a contained disk burst and was fragmented into multiple pieces. After the disassembly and examination of the engine was completed, various turbine parts were sent to the Rolls-Royce Material Laboratory for metallurgical examination. The parts were subjected to scanning electron microscope (SEM) examination to determine the fracture origin(s), fracture type, and any anomalies at the fracture origin as well as energy dispersive x-ray spectroscopy (EDS) and hardness measurements to determine if the parts were manufactured in accordance with the drawing specifications. All the turbine parts examined met the drawing specification with no material anomalies identified and all fractured parts, except for the power turbine stage 3 turbine wheel, fractured due to overload. The power turbine stage 3 turbine wheel was completely corn-cobbed with almost all the airfoils fractured transversely across the platform; one airfoil exhibited features indicative of a high cycle fatigue (HCF) fracture with initiation at the trailing edge and progressing forward approximately a half inch before the blade ultimately failed in overload. The rest of the power turbine stage 3 turbine wheel fracture surfaces showed indications consistent with overload. No anomalies were found in the material composition of the part, and no anomalies were found at the fracture origin. According to the maintenance records, the power turbine stage 3 turbine wheel was visually and fluorescent penetrant inspected at the last engine overhaul, approximately 45.5 hours and 99 cycles, prior to the accident flight. The power turbine stage 3 wheel crack was most likely not present during the last inspection, 99 cycles prior to the event. HCF is typically categorized as high frequency loading usually associated with resonance/natural frequency of the material driven by low energy levels or forced vibration at higher energy levels. Striations are markings found on cyclic fractures that correlate to alternating stress cycles where one striation typically equates to a single stress cycle. HCF striation counting to determine the length of time the crack propagated before failure is difficult because striations are hard to resolve (count) due to the low alternating stress condition and high number of stress cycles. Low cycle fatigue (LCF), on the other hand, is generally mission driven and occurs at lower frequencies and higher stresses, where the stress cycles correlate to a flight cycle. LCF striations are easier to observe and count to estimate of crack propagation time. One feature of HCF failures is that HCF striations will typically exceed the known mission flight cycles; therefore, it was unlikely that the crack was present during the last inspection, 99 flight cycles prior to the event. The engine’s electronic control unit (ECU) was downloaded and the engine operating and fault data showed that the first indications of an engine problem were a reduction of power turbine speed and torque. Several seconds later, a drop in fuel flow and gas producer turbine speed were observed, indicating a problem with the gas turbine section. The initial engine failure was due to the observed HCF fracture of a single power turbine stage 3 turbine wheel blade and the subsequent loss of the remaining blades in overload followed by the gas producer stage 1 turbine wheel burst. The most likely cause of the gas producer stage 1 turbine wheel burst was an overspeed event resulting in stresses that exceeded the material’s capability. A rotor disk burst is typically caused by either a material anomaly, such as a crack, that compromises the material design or an overspeed condition where the centrifugal loads exceed the capability of the material to withstand those stresses. Examination of all recovered gas producer stage 1 turbine wheel fragments exhibited overload features with no observed material anomalies. There were no overspeed events recorded on the ECU. The turbine-to-compressor coupling shaft connects the gas producer turbine rotor, which is located at the back of the engine, to the compressor rotor and accessory gearbox towards the front of the engine. The gas producer turbine speed sensor is located in the accessory gearbox. The accessory gearbox gears, the turbine-to-compressor coupling, and the compressor impeller were all in good condition and showed no signs of damage, which indicated the whole gas producer gear train did not overspeed. The gas producer stage 1 wheel most likely decoupled from the gas producer stage 2 wheel, which allowed it to overspeed while the remainder of the gas producer gear train continued to drive the accessory gearbox and compressor impeller. With the loss of the gas producer turbine stage 1 wheel, the Ng spool would decelerate. The drop in gas producer speed was observed in the engine control unit data shortly after the observed power turbine speed and torque reduction. The power turbine and gas producer turbine rotors are not mechanically coupled, but the power turbine bearing support provides a load path between the two spools. The power turbine bearing support provides support for the front (stage 2 turbine wheel) of the gas producer turbine rotor by the No. 7 roller bearing and the rear (stage 3 turbine wheel) of the power turbine rotor by the No. 6 roller bearing. The gas producer stage 3 turbine wheel failure created a sufficient imbalance where the power turbine rotor rotated “off-center’ or in an eccentric motion. This results in contact and rub which imparted high radial and axial loading from the power turbine rotor to the power turbine bearing support. Indications of “off-center” rotation of the power turbine rotor included: 1) stage 3 turbine wheel and the power turbine support contact with the power turbine labyrinth seal, 2) 360° circumferential rub on the power turbine shaft and the No. 6 bearing inner race land, and 3) rub on the power turbine bearing support front face outer diameter with corresponding contact on the inner diameter of the power turbine rotating labyrinth seal. The high radial and axial loading imparted through the power turbine bearing support to the gas producer rotor resulted in static hardware contact and rub on the gas producer stage 2 turbine wheel which caused increased drag and a speed mismatch between the gas producer stage 1 and stage 2 wheels. The gas producer stage 1 and stage 2 wheels are held together by a high compressive preload on the tie bolt that keeps the curvic coupling teeth properly aligned and loaded so that the gas producer transmits the torque at the same rotation speed. Since the gas producer wheels rotate at the same speed, there is minimum torsional/rotational load on the tie-rod during normal operation. However, when a mismatch occurs, the tie-rod will experience an unintended high torsion load that can cause the tie-rod to fracture. Metallurgical examination of the tie rod confirmed that it fractured due to torsional overload. The fracture of the tie-rod allowed the gas producer stage 1 and stage 2 wheels to separate and the gas producer stage 1 wheel to overspeed. All the stage 1 turbine wheel aft curvic coupling teeth and the corresponding stage 2 turbine wheel forward curvic coupling teeth were rotationally smeared over with very little tooth height remaining consistent with rotational speed mismatch between the two gas producer turbine wheels.

Verbatim NTSB analysis and probable cause from the NTSB dataset