HISTORY OF FLIGHT
On October 23, 2011, about 1600 eastern daylight time, a Boeing 747-451,registration number N661US, powered by four Pratt & Whitney PW4056 turbofan engines, operated by Delta Air Lines as flight 275, experienced a left-hand inboard engine (No. 2) low pressure turbine failure after takeoff from Detroit Metro Wayne Country International Airport (DTW), Detroit, Michigan. According to the flight crew statements, at about 5,000 feet above ground level following clean-up of the airplane, a large compressor stall and a muffled explosion was felt followed by the No. 2 engine rollback. A No. 2 engine fire warning indication illuminated and the No. 2 engine throttle was reduced to idle; however, the fire warning indication did not extinguish. The flight crew performed the engine fire procedure and the fire warning indication extinguished after discharging one fire bottle. The flight crew performed an air turn back and made a successful three-engine landing at DTW. Of the 373 passengers and 19 crewmembers on board the flight, no injuries were reported. The incident flight was a 14 CFR Part 121 international passenger flight from Detroit to Narita, Tokyo Japan (NRT). Day visual meteorological conditions prevailed at the time of the event, and an instrument flight rules flight plan was filed.
AIRCRAFT AND ENGINE DAMAGE
Examination of the airplane revealed impact punctures and holes to: 1) the left-hand wing, flaps, and ailerons, 2) the No. 2 engine strut fairings and panels; and 3) the left-hand horizontal stabilizer. Also noted were superficial scuff and scrape marks on the fuselage and under the left-hand wing. Initial examination of the engine revealed three burn-through holes on the No. 2 engine outboard core cowl and severe low pressure turbine (LPT) damage with many blades and vanes missing from multiple stages. Opening of the fan cowls, thrust reverser, and core cowls for the No. 2 engine revealed thermal distress and sooting from the fan exit rear case back to the rear turbine case, numerous exit hole penetrations in the rear turbine case, and a punctured No. 4 bearing compartment oil pressure supply line.
The No. 2 engine was removed from the airplane and shipped to the Delta Air Lines Engine Maintenance shop in Atlanta, Georgia where the Powerplant Group convened to perform a detailed examination of the engine. Disassembly of the engine revealed that: 1) no damage upstream of the stage 3 LPT was noted, 2) all the LPT blades were corn-cobbed (blades fractured transversely across the airfoil at the blade platform), 3) numerous vane clusters in each stage were either missing or heavily battered, 4) one stage 3 LPT outer transition duct segment and two stage 3 LPT vane clusters missing from their normally installed position, and 5) the majority of the stage 3 LPT outer transition duct segment rear seal plates were damaged, broken, and loose.
ADDITIONAL INFORMATION
Previous Stage 3 LPT Outer Transition duct Segment Failures
Pratt &Whitney (P&W) issued service bulletin (SB) PW4ENG 72-488 to address previously documented problems with the stage 3 LPT outer transition duct segment rear seal plates becoming loose and falling off, allowing hot gases to leak under the transition duct segment and causing them to distort and disengage from the rear turbine case. The SB provided repaired/modified procedures to incorporate new thicker rear seal plates and larger rivets to the existing stage 3 LPT outer transition duct segment configuration and also introduced a new configuration with an integral rear seal that eliminated the rear seal plates entirely. Review of the maintenance records (and later confirmed by physical inspection) revealed that the stage 3 LPT outer transition duct segments that were installed in the engine at the time of the LPT failure were the stage 3 LPT outer transition duct segments with the new thicker rear seal plates and larger rivets but not the integral rear seal configuration.
Including this event, this is the sixteenth known reported stage 3 outer transition duct segment(s) that either partially or fully disengaged from the rear turbine case. All the stage 3 outer transition duct segments liberations to date have been with the riveted seal plate configuration – none with the integral seal configuration. Fifty percent of the reported events were found during a shop visit and did not impact the operation of the engine, while the other half resulted in an in-flight shutdown or aborted takeoff. The first reported events occurred back in 1991, and prior to the release of SB 72-488, there had been 6 such events with 3 resulting in in-flight shutdowns. Of the sixteen known stage 3 outer transition duct segment(s) disengagements, the DAL event is the only one to result in turbine debris penetrating the rear turbine case and the only one where an undercowl fire was the result. Only one other event was investigated and documented sufficiently such that the information collected from that event could be useful in comparing the damage with the DAL event. That event occurred in August 2007 and was investigated by representatives of the P&W Flight Safety Office.
Based on the hardware damage, P&W determined that the cause of the LPT damage in the August 2007 event was related to the failure of the stage 3 outer transition duct segments rear riveted seal plates that allowed hot gas path air under the segments causing them to thermally distort and disengage from the rear turbine case. The liberated transition duct segments contacted the stage 3 vane clusters pushing them into the stage 3 blades resulting in liberation of stage 3 vane segments, fracturing of all the stage 3 blades, and causing downstream collateral damage. Comparing the photo documentation of the August 2007 event and the DAL event, many similarities in the type and extent of the damage to the stage 3 LPT and downstream hardware were observed.
Engine and Airplane Containment Requirements
Section 33.19 Durability
The engine containment standards are found in 14 CFR Part 33 Subpart B - Design and Construction; General, Section 33.19 ‘Durability’ and require engine manufacturers to design compressor and turbine rotor cases that must provide for the containment of damage from rotor blade failure. Examination of the engine revealed that the fan case sustained some bulging but no exit holes, penetrations, or uncontainments were noted. At the time that the PW4056 was certificated, FAA Advisory Circular (AC) 33-1B, dated April 22, 1970, provided guidance and acceptable means of compliance by which the engine manufactures test and certificate that the engine is in compliance with and rotor blade containment provisions in CFR Part 33. Under the substantiation criteria for rotor blade containment under AC 33-1B, it states that the engine is acceptable if, during the tests, the damage from rotor blade failures is contained by the engine, e.g., without causing significant rupture or hazardous distortions of the engine casing and the expulsion of blades through or beyond the edge of the engine case or shield. The multiple penetrations documented in the rear turbine case and the piercing of the No. 4 bearing pressure line shows that the PW4056, and more specifically the rear turbine case, did not meet the intent of Part 33.19(a) Amendment 9.
Section 33.75 Safety Analysis
Subpart E – Design and Construction; Turbine Aircraft Engines, Section 33.75 Safety Analysis requires that engine failures or malfunctions do not result in a hazardous or unsafe condition and that any probable malfunction or any probable single or multiple failure, or any probable operation of the engine will not cause the engine to catch fire, burst (penetrate its case), generate loads greater than those specified in Section 33.23(a), or lose the capability of being shut down. The engine and the left hand core cowl sustained fire damage as a consequence of the internal engine failure; thus the PW4056 did not meet the intent of 33.75 Amendment 6 requiring the engine not catch fire due to an engine failure or malfunction.