7O-FAB
BOMBARDIER CRJ700
On September 13, 2014, at about 06:02 UTC, Airways Bombardier (formerly Canadair) CRJ702ER Next Generation (NG), registration number 7O-FAB, powered by two General Electric (GE) CF34-8C5B1 turbofan engines, experienced a No. 1 (left) engine failure just after takeoff from the Hodeida International Airport (HOD), Yemen. The flight crew received a No. 1 engine fire warning indication along with high N2 vibrations and performed an inflight shutdown (IFSD) of the engine and made an uneventful air turn back (ATB) and landing at HOD. External examination of the No. 1 engine revealed a 26½ inches circumferential breach in the low pressure turbine case. Disassembly of the engine revealed a 10 inch burn-through hole in the combustion chamber inner liner. No anomalies were noted to any of the hardware upstream of the combustion chamber. Downstream of the combustion chamber, thermal distress and impact damage was noted throughout the high pressure and low pressure turbines. Debris that looked and felt like a brown soft talc was collected from the HPT aft shaft and HPT stage 2 blade retainers and sent to GE for analysis. Although the samples were similar, the HPT stage 2 blade debris was enriched with calcium sulfate (CaSO4). According to GE, the dust forms a calcium-magnesium-aluminum-silica (CMAS) that melts and infiltrates the structure on the thermal barrier coating (TBC) of the combustion liner surfaces causing the TBC to spall and accelerate thermal deterioration of the liner. Engine trend data for the event engine showed that starting in early May 2014 that there was an anomaly with the engine's interstage turbine temperature. GE recommended that the engine should be borescope inspected and to troubleshoot the temperatue indication system. Review of the engine logbook for the event engine found no mention of a borescope inspection or troubleshooting of the temperature indication system after May 2014. A review of other CF34-8C5/8E engines operating in a similar harsh enviroment found the event engine found had remained on wing and operated almost 2.5 longer than the average engine.
HISTORY OF FLIGHT On September 13, 2014, at about 06:02 UTC, a Felix Airways Bombardier (formerly Canadair) CRJ702ER Next Generation (NG), registration number 7O-FAB, powered by two General Electric (GE) CF34-8C5B1 turbofan engines, experienced a No. 1 (left) engine failure just after takeoff from the Hodeida International Airport (HOD), Yemen. The flight crew received a No. 1 engine fire warning indication along with high N2 vibrations and performed an inflight shutdown (IFSD) of the engine and made an uneventful air turn back (ATB) and landing at HOD. The operator reported that after landing, damage was observed to the turbine blades, holes were found in the low pressure turbine (LPT) case and the aft cowl, and that the No. 1 engine needed to be replaced. The event flight, flight No. 171, was a regularly scheduled flight from HOD to Sana´a International Airport (SAH), Yemen. ENGINE DAMAGE EXAMINATION External examination of the No. 1 engine revealed a breach in the LPT case which was located just aft of the LPT case forward flange and measured approximately 26½ inches circumferential from about the 6:30 - 11:30 o'clock position. The edges of the LPT case skin at the breach location showed damage consistent with thermal/melting distress. The only thermal distress to the outside of the engine was on the left side of the engine in the vicinity of the LPT case breach; it consisted primarily of melted: 1) anti-chaffing sleeves on the igniter cables, 2) electrical leads for the operability bleed valve (OBV), 3) interstage turbine temperature (ITT) probes, and 4) the No. 5 bearing oil supply tube. Disassembly of the engine revealed a 10 inch burn-through hole (material was consumed and missing) in the combustion chamber inner liner panels Nos. 2 and 3. The burn-through spanned circumferentially from the 10:30 - 00:30 o'clock position aft looking forward. The combustion chamber outer liners exhibited multiple linear axial cracks within the panels and emanating from the cooling holes, and thermal damage consisting of burned and melted material. The most significant thermal damage was to outer liner panels Nos. 2 and 3 and coincident with the inner liner burn-through location. No anomalies were noted to any of the hardware upstream of the combustion chamber. Downstream of the combustion chamber, thermal distress and impact damage was noted throughout the high pressure and low pressure turbines. TEST AND RESEARCH Debris that looked and felt like a brown soft talc was collected from the HPT aft shaft and HPT stage 2 blade retainers and sent to GE for analysis. According to GE, the debris samples were consistent with dust from the Middle East and the chemistry and particle size distribution of the two samples were almost identical. The only difference was that the dust from the HPT stage 2 blade debris contained no carbonate minerals and instead was enriched with calcium sulfate (CaSO4). GE suggested that the HPT stage 2 blade dust was exposed to sulfur dioxide (SO2) as a by-product of the combustion process that reacted with the carbonates to form CaSO4. According to GE , the dust forms a calcium-magnesium-aluminum-silica (CMAS), which melts on the thermal barrier coating (TBC) of the combustion liner surfaces. The molten CMAS infiltrates the structure of the TBC and contributes to early spallation of the TBC, which then causes accelerated thermal deterioration of the liner and early removal. The fuel nozzles were flowed tested to access their spray quality and to determine if there were any anomalies that would account for the thermal damage observed in the combustion chamber. Although the fuel nozzles in the as-received condition did not pass all the flow requirements of a new or overhaul fuel nozzle, no significant fuel streaks or spray angle deviations were observed and according to GE, the overall condition was typical of in-service units. The high pressure turbine (HPT) stage 1 and 2 blades showed significant thermal distress that included melting, consumed material, and missing coating. Metallurgical examination of sample HPT stage 1 and 2 blades, stages closest to the combustion chamber, revealed that those parts were subjected to temperatures above their intended maximum design point and material capability. Testing of both the upper and lower ITT harness assembles indicated that both harnesses were not capable of accurate, reliable, and stable temperature readings in the post-event state. Based on the thermal damage to the both harnesses, the manufacturer of the harnesses could not positively determine whether the erroneous temperature values recorded during the post-event testing were due to issues present before the engine fire event or were caused by the engine fire. ADDITIONAL INFORMATION A review of the event engine trend data showed that starting in early May 2014 that there was a decrease in ITT (an increase in exhaust gas temperature margin) indicating that the engine was running cooler. The data also showed a slight decrease in fuel flow and slight increase in core speed in comparison to a baseline CF34-8C5B1 engine. Felix Airways engineers noticed a significant difference in ITT between the engine event (installed in position No. 1) and the sister engine (installed in position No. 2 of the same airplane) with the sister engine between 20°C to 50°C higher than the event engine. Felix Airways provided that information to GE and requested assitance in what actions would be appropriate to take to address the higher observed temepature in the sister engine. GE reviewed the data and was more concerned with temperature of the event engine being lower instead of the sister engine temperature being higher. GE believed that the low temperature values on the event engine were related to an indication or sensing issue. GE suggested that both engines should be borescope inspected to determine the serviceablity of the engines and if deemed serviceable that the temperatue indication be troubleshooted. Review of the engine logbook for the event engine found no mention of a borescope inspected or troubleshooting of the temperature indication system after May 2014; therefore it could not be determined whether or not those actions were performed. GE conducted a review of removal rates for CF34-8C5/8E engines operating in what was defined to be "harsh and non-harsh" environments at the request of the Safety Board. Engines operating in a "harsh" environment were defined as "Engine[s] that operate for more than 50 percent of their departures within [long list of countries]…, or Yemen" with all other operations are defined as "non-harsh". The data showed that a large portion of the fleet that operate in a "harsh" environment have a shop visit between 3,000 – 3,500 cycles range and for those engines operating in a "non-harsh" environment have a shop visit near the 9,000 – 10,000 cycles range. At the time of the event, the engine had accumulated 8,010 flight hours since last shop visit and 9,347 cycles since last shop visit, indicating that the event engine was on-wing almost 2.5 times longer than the average engine in the "harsh" environment before removal. The calculated average flight leg was approximately 51 minutes.
The probable cause of the undercowl engine fire was the burn-through of the combustion chamber inner liner from the accumulation of dust that became molten during operating and infiltrated the combustion chamber thermal barrier coating causing it to spall off and accelerate normal thermal distress. The burn-though of the combustion chamber caused downstream hardware to experience higher than designed temperatures that resulted in mechanical and thermal distress to the point that the low pressure turbine case began to melt. Contributing to the combustion chamber failure was the operators' failure to do recommended borescope inspections that would have detected the combustion chamber distress and hence the removal of the engine before it was allowed to progress to the point that the combustion and turbine sections of the engine failed.
Source: NTSB Aviation Accident Database
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