Portland, OR, USA
N304FE
MCDONNELL DOUGLAS MD-10-30F
On February 13, 2012, a FedEx McDonnell Douglas (Boeing) MD10-30, registration number N304FE, equipped with three General Electric (GE) CF6-50 turbofans engines, performed a rejected take-off due to a failure of the No. 1 (left) engine. No injuries were reported. The incident flight was a regularly scheduled domestic cargo flight, operated under the provisions of Title 14 CFR Part 121, from Portland to Oakland, California. Examination of the airplane revealed no damage or penetration of the cowlings and no damage to the airplane; however, examination of the engine revealed three penetration holes and tears in low pressure turbine case and a single hole located in the turbine rear frame. Disassembly of the incident engine found that the fan mid shaft had fractured and the aft section of the fan mid shaft exhibited a spiral crack angled 45° to the longitudinal axis that measured about 69.5-inches in total length (about 3½ revolutions around the shaft) and extended about 20.5-inches axially in length. Metallurgical examination of the fan mid shaft found that it had fractured entirely in an intergranular mode. The initiation of the spiral fracture was comprised of numerous discolored, thumbnail shaped cracks initiating from the inner diameter surface of the fan mid shaft, some of which had progressed through the fan mid shaft thickness. The morphology of the cracks was consistent with a stress corrosion cracking mechanisms typical of high strength steel such as the fan mid shaft parent material. Localized areas of corrosion pitting, scaling, and loss of the anti-corrosion coating were observed throughout the inner diameter of the fan mid shaft. Examination of the debris within the shaft identified it as an ester based engine oil residue, typical of the kind of oil used in turbine engines. The stress corrosion cracking observed on the fan mid shaft was caused primarily due to the breakdown of a synthetic oil product inside the FMS during both storage and engine operation. The exact source or mechanism by which the oil entered the dry cavity between the fan mid shaft and the center vent tube is unknown.
HISTORY OF FLIGHT On February 13, 2012, a FedEx McDonnell Douglas (Boeing) MD10-30, registration number N304FE, equipped with three General Electric (GE) CF6-50 turbofan engines, encountered a No. 1 (left) engine stall, high vibration, and high exhaust gas temperature during the initial takeoff roll at about 60 knots at the Portland International Airport (PDX) Portland Oregon. The flight crew performed a rejected take-off and the fire handle was pulled. There were no reports of the fire extinguishing bottles having been discharged. The airplane taxied back to the gate and a normal deplaning was conducted. No injuries were reported. According to FedEx, visual examination of the airplane revealed no damage or penetration of the cowlings and no damage to the airplane. The incident flight was a regularly scheduled domestic cargo flight, operated under the provisions of Title 14 CFR Part 121, from Portland to Oakland, California. ENGINE DAMAGE Initial Visual Examination The engine was removed from the airplane and shipped to the FedEx maintenance facility in Memphis, Tennessee for initial visual examination. The Powerplant group comprised of members from GE, FedEx, and the National Transportation Safety Board (NTSB) convened in Memphis on February 23, 2012 to commence the visual examination of No. 1 engine, engine serial number (ESN) 517-558, and completed its work the following day. Examination of the engine revealed no signs of an undercowl fire or fluid leaks. When the fan was rotated, the low pressure turbine (LPT) rotor did not rotate along with it and when an attempt was made to rotate just the LPT rotor, it was seized. Three penetration holes and tears, located at the 4:00, 6:00, and 11:30 o’clock positions, were observed in low pressure turbine case, and a single hole, located at the 1:00 o’clock position, was observed in the turbine rear frame. Engine Disassembly and Examination The engine was shipped to the GE Celma facility in Petrópolis, Rio Janeiro, Brazil for a detailed teardown and examination. The Powerplant Group reconvened at the GE Celma facility on March 13, 2012 to commence the examination of the engine and completed its work on March 20, 2012. Disassembly of the engine revealed that the fan mid shaft, the center vent tube, and the high pressure compressor air duct were all fractured. The fan mid shaft was fractured about 18.5-inches aft of the forward end into two sections. The aft section exhibited a spiral crack that measured about 69.5-inches in total length (about 3½ revolutions around the shaft) and extended from the torsional fracture face aft about 20.5-inches axially in length. Over its entire length, the crack appearance changed from tight near the torsional fracture to wide, gapped, jagged, and offset in the middle part of the crack, then tight again near the aft end. The center vent tube was fractured into two pieces and was twisted and collapsed onto itself in the area of the fracture. The forward parts of the fan mid shaft and center vent tube remained attached to the fan forward shaft and the aft portions of those parts remained within the engine core. The outer surface of the center vent tube exhibited dried debris at several different locations. The high pressure compressor air duct was fractured into two major pieces (2 other smaller pieces were broken out of the larger pieces), with the forward part attached to the stage 2 high pressure compressor disk and the remaining aft parts resting on the fan mid shaft. In general, the high pressure compressor air duct exhibited circumferential wear, damage, and heat discoloration. The fan mid shaft, the center vent tube, and the high pressure compressor air duct, along with several other components/parts were sent to the GE facility in Cincinnati, Ohio, for metallurgical examination and evaluation. All the LPT blades exhibited blade tip fractures, leading and trailing edge airfoils impact damage, missing material, and some heat discoloration and all the stages 2, 3, and 4 LPT stator vanes exhibited outboard and inboard leading edge airfoil contact rub consistent with contact/clashing from the LPT blade tips. All the engine main line bearings were intact and in good condition except for the No. 7 bearing that was oil wetted but seized. Five of the No. 7 bearing roller elements were missing and all the remaining roller elements were heat discolored black and most were flattened. The No. 7 bearing cage was damaged and heat discolored but was intact. The inner race aft edge was damaged and material was rolled. TEST AND RESEARCH Metallurgical Examination Metallurgical examination of the fan mid shaft parent material composition and hardness measurements revealed it to be consistent with the required material composition and properly heat treated. The inner diameter of the fan mid shaft was coated with dry soot-like debris and some oil deposits but no residual liquid was present. Localized areas of corrosion pitting, scaling, and loss of the SermeTel® anti-corrosion coating was observed throughout the inner diameter of the fan mid shaft. Although not all the pits were measured, some pits were measured to be as deep as approximately 0.040-inches and some with pit diameters were up to approximately 0.067-inches. The spiral fracture was angled 45° to the longitudinal axis of the fan mid shaft, which is the primary stress plane of the fan mid shaft during engine operation. The spiral fracture is consistent with a torsional influence upon crack initiation and propagation. The spiral fracture was entirely intergranular. The initiation of the spiral fracture was comprised of numerous discolored, thumbnail shaped cracks initiating from the inner diameter surface of the fan mid shaft, some of which had progressed through the fan mid shaft thickness. The through cracks, similar to the spiral fracture, were torsionally oriented consistent with the primary stress direction of the fan mid shaft during engine operation, and originated in areas with corrosion pitting and deterioration of SermeTel® anti-corrosion coating. Further examination revealed that the through cracks were branched and had an intergranular appearance. The branched and intergranular morphology of the cracks was consistent with a stress corrosion cracking mechanisms typical of high strength steel such as the fan mid shaft parent material. Chemical composition analysis of the fracture surface revealed no causative species, such as chlorides or sulfides within the cracks; however, trace amounts of chloride was found around some of the corrosion pits. Debris found within the corrosion pits and along the inner diameter surface of the fan mid shaft and on the outer surface of the center vent tube was positively identified as an ester based engine oil residue, the kind of oil typical used in turbine engines. GE concluded that the stress corrosion cracking observed on the fan mid shaft was caused primarily due to the breakdown of a synthetic oil product inside the FMS during both storage and engine operation. The exact source or mechanism by which the oil entered the dry cavity between the fan mid shaft and the center vent tube is unknown. ADDITIONAL INFORMATION The LPT stage 2 and 3 interstage seals installed in the engine were the new configuration that were recommended by Safety Board Recommendation A-98-125, introduced by GE SB CF6-50 72-1268, and mandated by Airworthiness Directive 2005-26-06 to address safety concerns that when a fan mid shaft separation occurs it leads to a LPT stage 1 disk overspeed and uncontained failure. The pre-SB 72-1268 configuration LPT hardware allowed early contact and heavy rubs on the LPT stage 1 and LPT stage 2 disk spacer arms as the LPT rotor translates aft before the LPT blade and nozzle contact. This leads to rubs on the disk spacer arms that resulted in separation of the LPT rotor stage 1 disk and the disk to overspeed and then burst. The SB introduced new LPT stage 2 and 3 interstage seal hardware that have reduced axial length and a lower density honeycomb that would prevent heavy rubs on the LPT stage 1 and stage 2 disk spacer arms in the event of a FMS separation and ensure LPT blade airfoils will fragment by contact with the LPT nozzles, causing the LPT rotor to decelerate and prevent an overspeed and burst event. The meshing damage between the LPT blades and stators and the fact that the LPT stage 1 disk remained intact and still attached to the rest of the LPT rotor indicates that the SB 72-1268 configuration hardware performed as designed. Corrective Actions In an attempt to address the failure of the fan mid shaft due to oil migration inducted stress corrosion cracking, GE issued a series of corrective measures. On August 31, 2012, GE issued Service Bulletin (SB) 72-1323 providing instructions for a one-time borescope inspection of the fan mid shaft inner diameter to detect signs of corrosion pitting. The SB requests that all inspection results be reported to GE Product Support Engineering. In the January 2013 issue of the CF6 fleet highlights, which is monthly publication sent to all CF6 customers and that provides general information of interest or importance to operators, an article was published discussing this particular fan mid shaft failure event (no operator specified) and the rational for performing SB 72-1323.
The fracture of the fan mid shaft due to a stress corrosion cracking initiated by local regions of corrosion pitting, which resulted from oil migration within the shaft and its subsequent degradation.
Source: NTSB Aviation Accident Database
Aviation Accidents App
In-Depth Access to Aviation Accident Reports
