NTSB Factual Report

HISTORY OF THE FLIGHT

On March 2, 2012, at 1740 Hawaiian standard time, an Aerospatiale AS 355F1, N101MZ, experienced a partial loss of engine power and subsequent hard landing near Makakilo City, Hawaii. The helicopter was registered to TGR Helicopters, and operated by Rogers Helicopters, under the provisions of Title 14 Code of Federal Regulations, Part 133. The commercial pilot was not injured, and the helicopter was substantially damaged. Visual meteorological conditions prevailed; no flight plan had been filed.

The helicopter operator reported that during external load operations the helicopter experienced a loss of engine power from the right engine (No. 2) with an associated engine chip light. The pilot executed a running landing into a nearby landing zone. During the landing the helicopter encountered rough terrain, which resulted in deformation of the tail boom skin. A cursory examination of the No. 2 engine revealed separation of the exhaust and combustion tube from the engine. While under the oversight of a Federal Aviation Administration (FAA) inspector, the engine was removed from the airframe, placed in an engine crate, and shipped to Rolls-Royce for further examination.

HELICOPTER INFORMATION

The six-seat, twin-engine, conventionally configured helicopter, serial number 5045, was manufactured in 1980. It was powered by two Rolls-Royce Model A250-C20F 420-hp engines. Review of the helicopter’s maintenance records showed that the total airframe time, at the time of the accident, was 11,827 hours. The No. 1 engine, serial number CAE-840048, total time was 11,828 hours, 1,270 hours time since overhaul (TSO). The No. 2 engine, serial number CAE-832098, total time was 10,716 hours at the time of the accident, 1,014 hours TSO. A 30-hour airframe inspection was performed on February 26, 2012, at aircraft total time of 11,822.1 hours. The most recent 100-hour inspection was performed on January 8, 2012, at aircraft total time of 11,735.1 hours. The turbine assembly of the No. 2 engine was overhauled on January 27, 2005, and as of February 15, 2010, it had accumulated 1,276 hours. It was not determined if the accident engine’s turbine assembly had been involved in a one engine inoperative (OEI) event where it remained the operating engine since the last turbine section overhaul.

A Rolls-Royce Commercial Engine Bulletin, Engine, Turbine Assembly – Steady-State Operation Avoidance Range Limit, CEB A-1400, originally issued on December 22, 2006, with revision 3 issued on January 19, 2009, addressed engine N2 speed avoidance range for all Model 250 Series II engines. Depending on the exact model of the engine, the N2 avoidance range can be from 75% to 95% N2 steady state operation. The 3rd stage turbine wheel (T3), part number 23065833, is affected by this bulletin. Specific for all Eurocopter AS355 models, the P/N 23065833 turbine wheel is to be replaced at the next turbine overhaul. Prior to the overhaul if a one engine inoperable (OEI) condition is experienced by the helicopter, the T3 turbine wheel must be replaced on the engine that remained operating.

TESTS & RESEARCH

On July 10, 2012, the Allison (Rolls-Royce) 250-C20F, SN CAE-832098, was examined at the Rolls-Royce Corporation, Indianapolis, Indiana, under the supervision of an FAA inspector from the Indianapolis Flight Standards District Office (FSDO). No damage was observed on the compressor assembly, the compressor rotated freely, and was not disassembled. The accessory gearbox was turned by hand and continuity was noted between all accessory drive pads and the power turbine output drive. The combustion can liner appeared normal. The right-hand air discharge tube was observed pulled back approximately 1/4 inch form the combustion case arm. A small amount of 3rd stage turbine blade and vane material was identified in the exhaust collector. The 1st stage turbine nozzle assembly exhibited heat distress to one vane trailing edge. The 1st and 2nd stage turbine wheels were dirty but otherwise unremarkable. The 2nd stage turbine nozzle assembly was unremarkable. The 1st stage turbine nozzle shield was unremarkable. The 3rd stage turbine nozzle showed impact damage and missing vane material from all trailing edge vane airfoils. The 3rd stage turbine wheel exhibited one liberated blade with the fracture near the root. A 360-degree tip rub was noted in the 3rd stage turbine nozzle blade track as well as 360-degree leading edge rub indications were noted at the hub area. The 4th stage turbine nozzle showed 360° rotational scoring on the 3rd stage turbine wheel blade path and the 4th stage turbine wheel blade path. The leading edge flange was fractured as a result of the 3rd stage turbine wheel contact at the 6- to 8-o’clock positions. There was 360-degree rub indications noted on the stationary seal. There was trailing edge damage noted on (10) 4th stage turbine vanes. The 4th stage turbine wheel revealed (1) fractured blade at the tip/shroud area approximately 1/4 inch, (5) of the blades have leading edge damage and (1) leading edge shroud knife was fractured approximately 1/4 inch, 360-degree tip rub was noted to the 4th stage turbine wheel blade track. The turbine to compressor coupling was intact with moderate rub on the outer surface of the turbine end. Residual oil was found throughout the system. The No. 5 bearing retaining ring was found disengaged with the turbine section was removed. No damage was noted to the No. 5 bearing or bearing cavity. The No. 6 bearing was damaged due to heat distress. Residual fuel was found in the fuel system supply lines. Compressor bleed valve was intact with no apparent damage. The Fuel Control Unit (FCU) was found set in the minimum flow position.

The 3rd state turbine wheel assembly was shipped to the NTSB Materials Laboratory for further examination.

NTSB Materials Laboratory Examination Results

One blade was fractured near the platform surface. The third stage wheel was marked with p/n 23065833 in the hub area and serial number X53635 on the convex faces of successive blades. The wheel had 30 blades with integral shrouds connecting packets of 6 blades each. The fractured blade was the last (6th) blade in a packet when counted as the wheel rotates.

Magnified optical examination of the fracture revealed a relatively smooth, glossy dark fracture extending from the trailing edge though about 2/3 of the blade chord. Microscopic fracture markings and arrest lines within the dark region were indicative of fatigue progression from the trailing edge area forward. Closer view uncovered a single fatigue initiation site on the concave surface immediately adjacent to the trailing edge radius.

The forward third of the fracture was roughly textured with dendritic structures and a dark golden straw hue consistent with overstress separation in a cast material.

The fatigue origin was located in the airfoil about 0.045 inch to 0.050 inch above the platform surface. No mechanical damage was noted at the origin location. From the origin the fatigue propagated forward and slightly radially inboard intersecting the platform surface before turn radially outboard back into the airfoil. In total the fatigue penetrated about 0.55 inch.

The fractured blade and several adjacent blades were abrasively saw cut from the wheel, cleaned in acetone with ultrasonic agitation, and mounted for scanning electron microscope examination (SEM). Closer examinations uncovered highly oxidized surfaces with degraded fatigue striations.

NTSB Probable Cause

A partial loss of engine power while maneuvering over rough/uneven terrain that was precipitated by the failure of a 3rd stage turbine blade as a result of high-cycle fatigue.