LOS ANGELES, CA, USA
N87WC
Bell 206L-1
The helicopter lost engine power while on approach to a rooftop helipad, approximately 50 feet high and 50 feet short of the building. The pilot landed the helicopter on the roof, severing the tail boom at the horizontal stabilizer. The Safety Board conducted a teardown examination of the engine. Disassembly of the gas producer turbine support from the power turbine support revealed a fractured first stage turbine wheel, which displayed a nine airfoil chordal segment of the rim separated from the larger portion of the wheel. According to the Allison metallurgist, both the first and second stage turbine wheel blades exhibited evidence of operation of temperatures in excess of 2,200 degrees Fahrenheit. Both primary and secondary rim crack fractures exhibited evidence of exposure to extreme temperatures during the engine start sequence. No evidence of engine oil fire was observed on the gas producer turbine components. Chemical analysis of the affected gas producer turbine components indicated conformance to the specified material types. The Allison metallurgist reported that an overtemperature event occurred as a result of a hot start when both wheels were not at operating temperature; an overtemperature event occurring when the engine was at operating temperatures would have melted the first stage turbine wheel blades. The size of the gamma prime at the altered locations indicated the event occurred approximately within the last 10 hours of engine operation. No hot start or overtemperature event was reported to the maintenance facility.
HISTORY OF FLIGHT On April 20, 1998, at 2130 hours Pacific daylight time, a Bell 206L-1, N87WC, lost engine power while on approach to a rooftop helipad in downtown Los Angeles, California. The aircraft sustained substantial damage and the airline transport rated pilot and one crewmember, the sole occupants, were not injured. Night visual meteorological conditions prevailed at the time of the accident and a company VFR flight plan was filed. The helicopter was operated by Helinet under 14 CFR Part 135 as a nonscheduled domestic cargo operation hauling bank checks. The accident leg of the flight originated at the Ontario airport at 2110. The pilot reported that he was on approach to the rooftop heliport. Approximately 50 feet high and short of the east rim of the building, he lowered the collective and raised the nose slightly to slow his approach speed. He then heard a loud bang and felt a slight jolt. The pilot said that at that time the red overtemp light illuminated. He then pushed the nose over and fully lowered the collective in an attempt to stretch the glide to clear the building's edge and the large air conditioning units on the rooftop. The pilot stated that his sink rate had increased and he had a lot of forward airspeed, so he raised the collective and applied aft cyclic. The stinger and the aft portion of the skids then struck the roof surface. The helicopter bounced hard several times, spreading the landing gear, and severing the tail boom at the horizontal stabilizer. The helicopter came to rest in an upright position on the roof. DAMAGE TO AIRCRAFT There was no evidence of a postimpact fire. A section of the top and left side of the engine cowling showed localized scorching. A security officer who witnessed the arrival of the helicopter reported noting fire from the top of the aircraft prior to its impact on the roof top, but indicated that the fire had extinguished about the time the helicopter came to rest. A few small pieces of blade or vane material were noted laying in the bottom of the engine bay and were removed for further analysis. ENGINE EXAMINATION The Safety Board conducted a teardown examination of the engine on April 28, 1998, at National Airmotive in Long Beach, California. The gas producer support exhibited split-line separation between the gas producer support and the power turbine support. There was cracking noted on both sides of the split-line. The Gas Producer Support exhibited expansion associated with the area of cracking and also exhibited an associated axial crack and localized separation of the gas producer support. The cracking and deformation could be seen as a result of the expansion of the internal energy absorbing ring associated with the first stage turbine wheel plane. Disassembly of the gas producer turbine support from the power turbine support revealed a fractured first stage turbine wheel. The turbine wheel exhibited a nine airfoil chordal segment of the rim separated from the larger portion of the wheel. Affected gas producer turbine components from the engine were sent to Rolls Royce-Allison for a complete metallurgical evaluation. A copy of Allison's report is appended to this file. The parts and the report from Allison were also sent to the Safety Board's Materials Laboratory for review. Testing of the fuel system components, including the fuel control, governor, fuel pump and fuel nozzle, did not render any anomalies. A Barfield check of the TOT indicating system did not reveal any discrepancies. The lead-acid battery was tested and displayed an acceptable charge. The ignition module was tested and did not reveal any anomalies. METALLURGICAL EXAMINATION According to the Allison metallurgist, the failure of the first stage turbine wheel was due to accelerated creep crack growth, resulting from exposure to extremely high temperatures during one or more hot starts. Both the first and second stage turbine wheel blades exhibited microstructural evidence of operation at temperatures in excess of 2,200 degrees Fahrenheit. Both primary and secondary rim crack fractures exhibited fractographic evidence of exposure to extreme temperatures during the engine start sequence. Fluorescent Penetrant Inspection (FPI) of the first stage turbine wheel indicated 18 trailing edge rim crack indications and 3 leading edge rim crack indications on the failed first stage turbine wheel. No evidence of engine oil fire was observed on the gas producer turbine components. Chemical analysis of the affected gas producer turbine components indicated conformance to the specified material types. The gamma prime dissolution observed on both the first stage turbine wheel blade and the second stage turbine wheel blade indicated an overtemperature event occurred as a result of a hot start when both wheels were not at operating temperature. During the engine start sequence, the temperature drop between the two turbine wheels is very small; however, when the engine is at operating temperatures, the temperature drop between the first and second stage turbine wheels is significant. An overtemperature event occurring when the engine was at operating temperatures would have melted the first stage turbine wheel blades. The size of the gamma prime at the altered locations indicated the event occurred approximately within the last 10 hours of engine operation. Gamma prime in solution will reprecipitate when exposed to normal engine operating temperatures if given enough time. RESEARCH According to the aircraft discrepancy logs provided by the operator, there was no record of a hot start since January 25, 1996. The operator receives its maintenance from a separate facility. The chief mechanic was interviewed and reported that since the hot start in 1996, no overtemperature event was reported to him personally or to his facility. Examination of the maintenance facility's records did not reveal any anomalies or uncorrected defects. The records show that the last maintenance service rendered to the helicopter before the accident was a 100/200 hour inspection, which was performed on February 5, 1998, at which time the aircraft was determined to be in an airworthy condition. The helicopter accumulated approximately 65 hours between receiving the maintenance and the accident. The Allison Operation and Maintenance Manual for the 250-C28 Series engines was reviewed and a copy of the relevant portions is appended to this file. It states that there is no time limit allowed for any temperature over 1,830 degrees Fahrenheit, and if that temperature is reached, the maintenance action is to remove the turbine for heavy maintenance or overhaul.
An overtemp of the engine during start within 10 hours of the accident, and the failure to report the event to the maintenance facility.
Source: NTSB Aviation Accident Database
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