Ontario, CA, USA
N363CM
BOEING 767-338
The low pressure turbine stage 5 nozzle segment No. 3 was the only item identified that had any indication of primary fatigue, and no damage or distress was found upstream (forward) of the low pressure turbine that would have accounted for the damage observed throughout the low pressure turbine, the most likely initiating event was the fatigue fracture and liberating of the two missing stage 5 nozzle airfoils. The low pressure turbine stage 5 blades are located directly downstream (aft) of the stage 5 nozzle; thus, the loss of the two stage 5 nozzle airfoils would have traveled downstream in the airflow direction contacting and damaging the stage 5 blades resulting in stage 5 blade airfoil fractures. The damage inflicted on the stage 5 blades by the ingestion of the stage 5 nozzle airfoils would have created an imbalance in the low pressure turbine rotor. This low pressure turbine imbalance was confirmed by the flight data recorder data. Shortly after takeoff during climb the low pressure turbine vibration started to climb and within a few seconds reached the value of 5 cockpit units which is the maximum value that can be recorded; a cockpit unit is a dimensionless scaler unit, a magnitude with no unit of measure attached. GE performed a rotor imbalance and deflection analysis to determine the amount of radial deflection each stage of the low pressure turbine would be anticipated to experience based on the condition of the hardware as documented during the engine exam. Since the exact condition of the low pressure turbine hardware for any point in time during the event was not known the analysis does not provide the exact vibration levels or deflections at the time of the event or during the engine deceleration but instead a general assessment. The results of the analysis showed that the amount of low pressure turbine rotor expected radial deflection was several times greater than the nominal running blade radial clearances at takeoff for each of the low pressure turbine rotor stages. This is consistent with all the observed gouging and heavy wear down to the backing strip of the honeycomb on the blade outer shroud segments and the accompanying loss of blade tips for each stage of the low pressure turbine. The analysis also predicted that the low pressure turbine stage 5 would be the stage that experienced the most deflection and that the deflection would be the greater not at takeoff/climb low pressure turbine rotor speed but as it decelerates from the accumulation of damage and the pilot’s action to shut down the engine. This would indicate that the overall damage observed throughout the low pressure turbine was initially caused by the loss of the low pressure turbine stage 5 nozzle airfoils due to fatigue and impacting the stage 5 blades fracturing them creating imbalance and deflection in the low pressure turbine rotor. The low pressure turbine rotor imbalance and deflection progressively increased due to the accumulation of additional stage 5 blade damage and the accompanying loss of rotor speed that eventually led the entire low pressure turbine rotor to lose radial blade clearance. This loss of radial blade clearance throughout the low pressure turbine resulted in blades contacting static structure, fracturing, and causing additional downstream damage. Since the flight data recorder does not record vibration levels or amplitudes above 5 cockpit units, any vibration values higher than 5 cockpit units is capped to 5 cockpit units, the exact vibration the right engine experienced was unknown; however, the results of the GE imbalance and deflection analysis indicated that the loads experience by the right engine would have been sufficient to fracture the oil supply tube and the turbine exhaust sleeve. Since neither of these items showed signs of a pre-existing anomalies, their failures were as a resulted of the high vibrational loads from the right engine during the failure sequence. The low-grade thermal/fire damage observed on the inside of the right engine core cowls and thrust reverser halves and on the outside of the right engine was due to the oil from the fractured oil supply line contacting hot engine cases and smoldering and igniting. The fuel system, forward of the left and right fuel manifold where the thermal/fire damage was most pronounced, was leak checked and no leaks were found; thus, the only source of a flammable fluid would have been the fractured oil supply line.
HISTORY OF FLIGHT On January 29, 2021, about 10:49 pacific standard time, a Boeing 767-300, registration number N363CM, operated by Aerotransportes Mas de Carga, S.A. de C.V. as flight MMA6853, and powered by two General Electric (GE) CF6-80C2-B6 turbofan engines, experienced a right (No. 2) engine fire during initial climb right after takeoff from Los Angeles International Airport (LAX), Los Angles, California. The flightcrew reported feeling vibrations in the airplane as the landing gear was cycled up, so a second cycling of the gear was attempted, and the vibration continued. Shortly after that, the flightcrew detected a strong odor consistent with burning and a right engine fire warning message displayed on the engine indicating and crew alerting system (EICAS) display. The flightcrew declared an emergency, performed the Quick Reference Handbook (QRH) engine fire procedures, which included shutting down the affected engine and discharging 1 fire suppression bottle, and diverted to the Ontario International Airport (ONT), Ontario, California for an uneventful single engine overweight landing with no reported injuries to any of the flightcrew members. The incident flight was a 14 Code of Federal Regulations (CFR) Part 129 cargo flight from LAX to the Mexico City International Airport (MEX), Mexico City, Mexico. ON-SCENE AIRPLANE AND ENGINE DAMAGE ASSESSMENT A Powerplant Group was formed and comprised of members from GE, Boeing, Federal Aviation Administration (FAA), and the National Transportation Safety Board (NTSB), and Aerotransportes Mas de Carga, S.A. de C.V. (referred to for the remainder of this report as Mas Air Cargo). Assisting in the on-scene part of the investigation were personnel from Jett Pro Line Maintenance stationed at ONT and ONT operations center. The NTSB did not travel to ONT to examine and document the airplane and engine damage. Instead, Jett Pro and ONT operations center personnel provided initial pictures and descriptions of the damage. On January 30, 2021, the day following the event, Boeing field representatives arrived on-site to photo document the damage along with providing written field notes. Several days later, a GE representative arrived on-site to assist in the removal of the engine and securing it into a transportation stand for shipping to the GE Evendale, Ohio facility for examining and disassembly. The on-scene examination of the airplane revealed thermal distress and fire damage to the inside of the right engine core cowl and thrust reverser. The entire outer portion of the right engine turbine exhaust nozzle was missing as was the aft portion of the inner sleeve; the forward portion of the inner sleeve remained attached to the engine. The right engine strut/pylon outdoor skirt in-line with the aft edge of what remained of the turbine exhaust nozzle exhibited gouging and scratch marks. The airplane exhibited some impact marks and dings on the underside of the right wing and flaps, on the vertical and horizontal stabilizer, and on the right cargo door; however, some of the impacts appeared old and could not be confirmed if all the damage observed was from this incident. The on-scene examination of the No. 2 engine revealed no signs of engine uncontainments but did show signs of low-grade thermal distress comprised of sooting and melted, consumed, and damaged cushion clamps, fire detector loop isolators, and electrical wire outer sheathing. Looking at the front of the engine, no damage was observed to the fan blades and the fan rotated freely by hand; the low pressure turbine rotated along with the fan. Looking through the engine exhaust: 1) all the low pressure turbine stage 5 blade roots were still installed in the disk and all the blades exhibited a combination of airfoil transverse fractures at various lengths and airfoil impact damage, tears, gouges, and missing airfoil material, 2) all the low pressure turbine stage 5 blade outer shroud segments were still present and exhibited a combination of gouging, missing material, heavy rub, and the honeycomb was worn down to the backing strip, 3) low pressure turbine stage 5 nozzles segments were present and exhibited a combination of trailing edge airfoil impact damage, gouges, tears, and missing material, and 4) the turbine rear frame exhibited a combination of multiple cracks, tears, openings, and punctures holes; no signs of low pressure turbine debris penetrating through the turbine rear frame skin was observed. The oil tank was low of oil and debris was noted on the magnetic chip detector. When the integrated drive generator was removed in preparation for transportation of the engine, a fractured oil supply line that runs from the aft side of accessory gearbox behind the integrated drive generator was observed (Photo 1). The No. 2 engine, engine serial number 695440, was removed and shipped to GE for disassembly and examination. Photo 1: Fractured Gearbox Oil Line DETAILED EXAMINATION OF THE NO. 2 ENGINE – ESN 695440 The Powerplant Group comprised of members from GE, FAA, Mas Air Cargo, and the NTSB convened at the GE Evendale facility from February 22-26, 2021, to perform a detailed examination of the incident engine, engine serial number 695440. During the pre-disassembly/induction borescope inspection: 1) no damage was noted to the combustor section, 2) all the high pressure turbine stages 1 blades were present, and no impact damage was noted but many exhibited thermal distress consistent with mid-span core burn-through, 3) first indications of event related turbine damage or distress was to the leading edge of the low pressure turbine stage 1 blades, and 4) damage was observed throughout all five stages of the low pressure turbine rotor. Each stage of the low pressure turbine exhibited combinations of the following damage to various degrees: 1) significant amount of leading and trailing edge blade and nozzle segment airfoil impact damage, material loss, and thermal distress except for the stage 1 nozzle segments that only exhibited minor thermal distress and minor trailing edge impact damage, 2) all blade roots were installed in their respective disk and all were fractured transversely across the airfoil, 3) blade outer shroud segments were gouged, the honeycomb worn down to the backing strip, and were thermally distressed, 4) disk rub and contact marks, 5) the rotating interstage air seal knife edges contact rub and material loss, 6) stationary honeycomb interstage air seal land trenching, gouging, material loss, and attachment bolt contact rub, and 7) consistent throughout each stage was evidence of contact between rotating and stationary hardware. With all the impact damage, contact rub, and thermal distress to each of the low pressure turbine stages observed, a few items were of particular interest. One stage 5 nozzle segment, labelled as the No. 3 segment, was missing two complete adjacent airfoils labeled airfoil Nos. 5 and 6; each stage 5 nozzle segment is comprised of 6 individual airfoils (Photo 2). No other nozzle segments from any other low pressure turbine stage were missing airfoils. Although all the low pressure turbine blades were fractured transversely across the airfoil at various lengths, none were fractured at or close to their platform and all the blade roots remained in the disk. The low pressure turbine case exhibited numerous outward impacts, bulges, tears, and a small hole in line with the stage 4 blades, none appeared to be a penetration hole where low pressure turbine debris passed through it. Photo 2: LPT Stage 5 Nozzle Segment Missing Airfoil METALLURGICAL EXAMINATION GE evaluated the material condition of the low pressure turbine hardware, the turbine rear frame, and the fractured oil supply line at their metallurgical laboratory in Evendale, Ohio while Boeing evaluated the material condition of the exhaust hardware, which included the turbine exhaust sleeve and the exhaust plug, at their Boeing Equipment Quality Analysis laboratory in Seattle Washington. The GE metallurgical analysis found the cracks and tears in the turbine rear frame and the turbine exhaust were due to overstress and no anomalies or pre-existing defects were noted that would have contributed to their condition. The fracture surfaces of the oil supply tube exhibited plastic deformation and a planar fracture consistent with secondary high amplitude fatigue/cyclic tensile loading. None of the low pressure turbine blade fracture surfaces showed signs of fatigue but features consistent with an overload failure. Signs consistent with high cycle fatigue growth were found on the fracture surfaces of airfoil Nos. 5 and 6 of the low pressure turbine stage 5 nozzle segment No. 3; the fatigue cracks propagated from both the leading and trailing edges of the missing airfoils (Photos 3 and 4). In the interpretable areas (not damaged area), GE found no material anomalies in the base material, aside from the areas of intergranular oxidation. Photo 3: Stage 5 Nozzle Airfoil Fatigue Fractures – Top View – Yellow Arrows Show Fatigue Propagation Photo 4: Fatigue Fractures – Airfoil View Low pressure turbine stage 5 nozzle segment No. 3, part number 9367M85P10, serial number HCM21194, had an extensive overhaul repair conducted in 2014 and has been in operational service for 10,232 hours and 4,068 cycles since the overhaul repair was performed. at the time of the event. Due to the leading and trailing edge impact damage on airfoils Nos. 5 and 6 of the low pressure turbine stage 5 nozzle segment No. 3 at the fracture locations, no determination could be positively made on whether a repair had been previously performed in that area, so no assessment of the repair conformity was made. The Boeing metallurgical analysis of the turbine exhaust sleeve revealed that all the fractures were composed entirely of ductile separation fracture mode, consistent with a single event, and no other anomalies were observed contributing to the fractures. LOW PRESSURE ROTOR DEFLECTION ANALYSIS GE performed a rotor imbalance and deflection analysis to determine the amount of radial deflection each stage of the low pressure turbine would be anticipated to experience based on the condition of the hardware as documented during the engine exam. Since the exact condition of the low pressure turbine hardware for any point in time during the event was not known and the vibration monitoring system only registers a maximum value of 5 cock pit units, the analysis does not provide the exact vibration levels or deflections at the time of the event or during the engine deceleration and may not represent the worst-case or maximum rotor imbalance or deflection (hence the highest engine vibration level) that the engine experienced during the incident flight. However, the analysis did provide a general assessment of the vibration levels and the amount of deflection the low pressure turbine was expected or predicted to experience after the event as the engine decelerated. The results of the analysis showed that the amount of expected radial deflection was several times greater than the nominal radial blade running clearances at takeoff for each of the low pressure turbine rotor stages. The observed gouging and heavy wear down to the backing strip of the honeycomb on the blade outer shroud segments for each stage of the low pressure turbine was consistent with the loss of radial blade clearance as predicted by the rotor imbalance and deflection analysis. The analysis also predicted that the low pressure turbine stage 5 would be the stage that experienced the most deflection and that the deflection would be the greater not at the takeoff/climb low pressure turbine rotor speed but as it decelerates from the accumulation of damage and the pilot’s action to shut down the engine.
The fatigue fracture and liberation of two airfoils from a low pressure turbine stage 5 nozzle segment that impacted and damaged the downstream low pressure stage 5 blades creating an initial imbalance load in the engine’s low pressure turbine rotor sufficient to allow all the low pressure turbine blades to lose radial blade clearance, contact static structure, and to fracture transversely across the airfoil. The progressive failure of the low pressure rotor caused an increasingly imbalanced load that eventually resulted in the fracture of the oil supply tube that allowed oil to contact hot engine parts and smolder and ignite resulting in the undercowl fire.
Source: NTSB Aviation Accident Database
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