NTSB CAROL · Event
Event DCA06FA033
Aircraft involved
Probable cause & findings
The rupture of a fatigue crack in the forward air seal tooth serration of the No. 3 engine low pressure turbine stage 1 disk that initiated due to an improper weld repair performed during original manufacture. The improper weld repair changed the material properties of the disk near the weld feature, resulting in an accelerated crack propagation rate that allowed the crack to grow to rupture before the next inspection interval. Contributing to the severity of the accident was the inability of the flight crew to shutdown the engine as a result of the damage caused by a liberated piece of the low pressure turbine stage 1 disk when it punctured the right wing.
Factual narrative
HISTORY OF FLIGHT
On April 4, 2006, at about 16:45 central standard time, FedEx Express flight 597, a McDonnell Douglas MD-10-10F, N386FE, experienced a No. 3 (right-hand) engine uncontained failure of the low pressure turbine (LPT) stage 1 disk while climbing through about FL260. About 8 minutes after the engine failure, and while the airplane was returning to Memphis International Airport (MEM), Memphis, Tennessee, the airplane experienced severe vibration that stopped when the entire LPT rotor separated from the engine. There were no injuries to the two flight crewmembers or one passenger onboard and the airplane received substantial damage. The flight was operating under the provisions of 14 Code of Federal Regulations (CFR) Part 121 from MEM to Seattle-Tacoma International Airport (SEA), Seattle, Washington. The captain was the flying pilot. The takeoff and initial climb phases of the flight were normal and uneventful. According to pilot interviews and recorded information, about 19 minutes after takeoff, as the airplane was climbing through FL260, a loud bang was heard and the airplane yawed to the right. An amber “X” appeared on the No. 3 engine instrumentation and the captain called for the Engine Fire and Severe Damage checklist. The first officer called for the passenger, a qualified MD-11 first officer who was in the courier compartment, to come to the cockpit to assist with the emergency. The first officer declared an emergency with air traffic control (ATC) and then completed the Severe Damage checklist, including shutting down the No. 3 engine and pulling its respective fire handle. The captain turned the airplane back towards MEM and the crew began dumping fuel. About eight minutes after the engine failure, the moderate airframe vibration that started when the engine failed, began to increase. This vibration intensified to a severe level over about a 6-minute period and then stopped abruptly about 14 minutes 38 seconds after the engine failure. In addition to the engine failure indications, the crew received No. 3 hydraulic pressure low, slat disagree, and right main gear unsafe warnings. The crew performed the checklists applicable to each of these warnings. The crew continued the descent and requested vectors for an extended final and visual approach to runway 36C at MEM. Due to the slat disagree warnings, the captain conducted a no slat, Flaps 22 approach with an approach reference speed (Vref) of 199 knots. After stopping on the runway, the aircraft rescue and firefighting (ARFF) personnel communicated that there was no fire present. As a result, the pilots shut down the No. 1 and 2 engines and started the auxiliary power unit while waiting for stairs to disembark. About 11 minutes after stopping, the ARFF personnel advised the flight crew that there was an increase in smoke around the No. 3 engine at which point, the captain ordered the crew to evacuate via the L1 (forward left) door slide. DAMAGE TO No. 3 ENGINE After landing, the LPT rotor was found to have separated from the No. 3 engine and a piece of the LPT stage 1 disk was lodged in the upper skin of the right wing. The LPT rotor wreckage was recovered from a field near Weiner, Arkansas. The flight data recorder (FDR) and cockpit voice recorder (CVR) data indicated that the LPT stage 1 disk failed over a forested area near Ravenden, Arkansas. An alert was issued for public assistance with finding the remainder of the disk, and a physical search was conducted of the forested area. No additional portions of the disk were recovered. There was no evidence of fire on the No. 3 engine, its respective cowling and thrust reverser, or on the right-hand wing. The aft engine mount beam and links were intact with no apparent cracks and were not distorted or twisted. The LPT case was fractured circumferentially and torn just aft of the LPT stage 1 nozzles where the LPT stage 1 disk would have normally been located. All the LPT case-to-turbine mid frame (TMF) attachment bolts were intact and the entire LPT case front flange remained attached to the rear flange of the TMF; however, the flange was fractured at three separate locations. Portions of the LPT case body, ranging from about 1-inch to 5-inches, remained with the front flange. Visual examination of the TMF did not reveal any damage, breaches, or signs of fire. With the aft end of the engine missing, the “C” sump housing, the LPT nozzle support, the static pressure balance seal, the LPT stage 1 interstage seal, and the LPT stage 1 nozzles were all exposed. All of the LPT stage 1 nozzles were intact, in place, and the airfoil trailing edges exhibited impact marks, tears, missing material, and were pushed forward. The inner ends of three nozzle segments, located at approximately the 2:30, 4:00 and 5:00 o’clock positions, were found disengaged from the LPT stage 1 interstage seal aft flange. The LPT stage 1 interstage honeycomb seal was intact and exhibited a deep trench along the middle of the seal down to the backing strip. The aft half of the honeycomb seal was also heavily worn, but not down to the backing strip. The rear flange of the LPT stage 1 interstage seal was distorted around the attachment bolts and exhibited circumferential contact wear and three radial impact marks located at approximately the 7:00, 8:30, and 10:30 o’clock positions. Damage at the 7:00 and 10:30 o’clock position consisted of a straight impact slash across the honeycomb material while the damage at the 8:30 o’clock position was a combination of a straight slash across the seal and aft flange, continuing into a spiral slash along the aft flange. The static pressure balance seal is comprised of an outer and inner seal land that works in conjunction with the rotating pressure balance seal attached to the LPT rotor to balance the pressure load of the HPT rotor and works with the No. 4 ball bearing to axially position the HPT rotor correctly within the engine. The outer seal land portion of the pressure balance seal was torn from the rest of the structure aft of the stiffener ring and exited the engine with the rest of the LPT rotor. This piece was recovered with the LPT rotor and turbine rear frame (TRF). The inner seal land remained attached to the support and was intact but was heavily distorted, flared outwards, compressed and exhibited material transfer, bluing, and heat distress. The inner seal land was also canted in relation to the engine’s longitudinal centerline – at the 6:00 o’clock position the seal was approximately 1.0 inch farther forward than at the 12:00 o’clock position. The “C” sump aft stationary air and oil seal is a single piece seal that works in conjunction with the rotating air and oil seal attached to the LPT rotor to prevent oil from leaking out of the bearing compartment and to provide pressure balancing throughout the bearing compartment. The oil-seal portion of the “C” sump seal is comprised of a “wind back” knife edge seal (forward most seal) and a nickel graphite seal land. The knife edges of the “wind back” seal were heavily damaged and flattened, and in some areas were completely rubbed away. Both oil-seal locations were heavy distorted, blued, missing, and exhibited heat distress. The airseal portion of the “C”-sump seal is comprised of a seal land and it was distorted, flattened forward, missing material, and exhibited bluing, material transfer, and heat distress. The outer race of the No. 6 bearing, which is located within the “C” sump, was fractured into several pieces, incomplete, and what remained was not oil wetted. Those pieces of the No. 6 bearing outer race that remained installed exhibited impact damage but no thermal distress or bearing roller material transfer. Fractured pieces of the No. 6 bearing cage were recovered loose within the bearing compartment and they exhibited heavy impact damage; however, no significant thermal distress or bluing was noted and the silver plating was still present. The LPT rotor (minus the stage 1 disk) and the TRF, along with various loose LPT blade and vane material, were recovered together in the same 3.5 foot deep crater near Weiner, Arkansas. The outer seal land of the static pressure balance seal was recovered approximately 190 yards from where the LPT rotor was found. The recovered LPT rotor was comprised of the stages 2, 3, and 4 disks, the LPT front shaft, No. 6 bearing inner race, the “C” sump aft air/oil seal, the aft portion of the fan mid shaft (FMS), the rotating pressure balance seal, the “D” sump air and oil seals, the LPT rear shaft, and the No. 7 bearing inner race. Some blades from each stage were missing and those that remained were fractured at various lengths. Forty-five degree shear lips, bluing, heat discoloration, and areas of melting/plastic flow were noted on the FMS fracture surfaces. The turbine rear frame (TRF) was comprised of a portion of the LPT turbine case (bottom half) still attached to the TRF outer case, the LPT stage 5 disk, the “D” sump housing, the center body, and exhaust nozzle.
DAMAGE TO AIRCRAFT
As a result of the engine failure, the airplane sustained numerous impact marks, holes, and punctures in the right-hand wing, the right-hand slats and flaps, the right-side of the fuselage, the No. 3 engine pylon, and the No. 3 main landing gear door. An airplane damage scatter pattern diagram was produced to map the approximate trajectory angles of debris as it was shed from the engine. The No. 3 engine pylon was intact, remained attached to the wing, and exhibited no holes or penetrations (thru holes). It did, however, exhibit scuffmarks and impact damage. Two sets of “train track” marks were noted on the outboard vertical skin of the pylon. The size, shape, and spacing of the “train track” marks were consistent with the size, shape, and spacing of the LPT stage 1 disk rim blade retaining fir tree slots. The pylon heat shield was severely impact damaged, torn, and the majority of it was missing from the pylon. The pylon fairing was intact, remained attached to the wing, and exhibited no damage except for the front face, which was torn exposing the inner structure. The right-hand side forward wing root fillet outer skin (fillet panel 154CB) exhibited an approximate 2 x 1 inch thru hole along with an approximate 1.5 x 0.25 inch scuffmark leading to it. This through hole was located approximately 10 degrees longitudinally and 5 degrees radially from the centerline of the engine and the LPT stage 1 disk reference plane. Removal of fillet panel 154CB revealed that one of the outermost vertical support hangers used to support the fillet panel brace was broken along with its mating fuselage attachment bracket. The right-hand main landing gear up-lock roller and attach fitting were missing from the landing gear door. No holes or penetrations were observed on the outside of the door. Fasteners that hold the gear up-lock and attach fitting to the landing gear door were missing and their respective holes were distorted. The structure around where the gear-up lock had been attached was ripped and torn rearwards. The majority of the right-hand wing damage was inboard of the No. 3 engine and concentrated radially from the No. 5 slat track to the pylon and longitudinally within a couple of feet aft of the LPT stage 1 disk reference plane; however, damage was also noted outboard as far as the No. 8 slat and rearward on the inboard flap hinge fairing. Approximately 16 feet outboard from the fuselage along the wing leading edge rivet line and 2.75 feet from the leading edge was a triangular shaped penetration hole in the top of the right wing that measured approximately 13 x 12 x 7 inches. The upper wing skin in the vicinity of the hole was torn and flared upwards and outwards. Protruding from the hole was a piece of the LPT stage 1 disk. After accessing the area, damage to the following electrical wires, hydraulic lines, and control cables was noted: No. 3 engine bleed duct severed; fire extinguishing agent supply line was severed; No. 3 hydraulic system main pressure line was severed; No. 3 hydraulic system main suction line was severed; droop leading edge mechanism (cable driven) was fractured; slat retraction pressure line was severed; slat extend pressure line was severed; hydraulic case line was crushed; No. 3 engine fire shutoff valve cable was severed (pylon fuel shut-off valve found in the OPEN position); No. 3 engine power off cable (throttle cable) was severed; right outboard slat follow-up retract cable was severed; the emergency shutoff cable for the No. 3 engine was severed; and an electrical bundle running through the damaged area was severed and exhibited cut wires. The front wing spar was intact and had no impact damage. The right wing No. 2 fuel tank was punctured in-line with the No. 2 slat. According to maintenance personnel, fuel was not noticed leaking from the tank until the airplane was being towed to the maintenance hangar. The outboard surface of the inboard flap hinge fairing exhibited three thru holes. A “train track” mark was also noted on the outboard surface of the flap hinge and it measured approximately 5 inches long. The underside of the right wing inboard of the No. 3 engine and the No. 2 slat exhibited a multitude of impact marks, scuffing of the skin, skin tears, holes and penetrations (thru holes). “Train track” marks were noted on the bottom of the right wing.
PERSONNEL INFORMATION
The captain, age 50, held an airline transport pilot certificate with airplane multiengine land and DC9, B727, DC10, and MD11 type ratings. The captain held a first-class medical certificate dated October 5, 2005. FedEx Express records indicated that the captain had accumulated 10,000 total flight hours, including 979 hours pilot in command in the MD-11 airplane. In the 90 days, 30 days, and 24 hours before the accident, the captain had flown 120, 45, and 1 hours, respectively. He received his last proficiency check on March 2, 2006. The first officer, age 45, held an airline transport pilot certificate with airplane multiengine land and MD-11 type rating. The first officer held a first-class medical certificate dated March 21, 2006. FedEx Express records indicated that the first officer had accumulated 6,900 total flight hours, including 333 hours in the MD-11 airplane. In the 90 days, 30 days, and 24 hours before the accident, the first officer had flown 98, 22, and 1 hours, respectively. He received his last proficiency check on December 20, 2005.
AIRPORT INFORMATION
MEM is located about 3 miles south of Memphis, Tennessee, and has an elevation of 341 feet mean sea level. The airport has four runways: 9/27, 18R/36L, 18C/36C, and 18L/36R. Runway 18C/36C is grooved concrete, and is 11,120 feet long and 150 feet wide.
FLIGHT RECORDERS
The airplane was equipped with a Honeywell solid-state flight data recorder model 980-4700. The recorder was in good condition and the data were extracted from it normally. The airplane was equipped with a 2-hour Allied Signal solid-state cockpit voice recorder model 980-6022-001. The CVR was in good condition and the audio information was extracted from it normally. A summary of the accident flight recording was produced.
ENGINE DISASSEMBLY
The No. 3 engine was removed from the airplane and examined at the General Electric (GE) Aviation’s Caledonian services facility in Prestwick, Ayrshire, Scotland, from May 3-7, 2006. The damage to the engine was documented and the forward part of the fan mid shaft, the high pressure turbine aft shaft, and the low pressure turbine case were removed for metallurgical analysis at the NTSB Materials Laboratory. The low pressure turbine (LPT) rotor was examined at GE Aviation’s facility in Evendale, Ohio, to measure and evaluate the bore diameters of the LPT disks for evidence of over speed. The disks were found to be heavily warped; consequently an accurate speed estimate based on the physical evidence alone could not be made.
METALLURGICAL EXAMINATION
LOW PRESSURE TURBINE STAGE 1 DISK Metallurgical examination of the fracture surfaces of the recovered LPT stage 1 disk revealed features consistent with a mixture of coarse and fine fatigue striations corresponding to propagation under spectrum load conditions. The fracture started in the forward seal tooth serration arm and grew in a plane perpendicular to the arm surface, then changed to a slant plane as it progressed through the rim, intersecting a blade slot bottom before the disk finally ruptured. Since the tip of the seal tooth was missing, the fatigue origin could not be identified. Microstructural sections through the seal tooth and arm revealed a fine dendritic structure in the radial direction with a boundary located at about the middle of the seal tooth height and a coarse columnar dendritric structure that had an axial direction in the lower half of the seal tooth that extended into the forward arm. These features were consistent with having been produced by at least two separate weld passes from different directions with the forward face having been welded from the forward side and the tooth tip from the radial direction.
OTHER ENGINE COMPONENTS
The fan midshaft had rubbing and heating damage consistent with the imbalance following the stage 1 LPT disk failure. The high pressure turbine rear shaft had fatigue, which initiated from an area of circumferential rubbing that was consistent with being secondary to the stage 1 LPT disk failure. The LPT case fracture surfaces in general were on slant angles and gray in color, with the features consistent with overstress.
MAINTENANCE HISTORY
According to maintenance records, the LPT stage 1 disk installed in the No. 3 engine was part number (PN) 9079M51P03, serial number (SN) MPOAP037. According to GE records, manufacture of disk SN MPOAP037 started in June 1988 at the GE facility in Evendale, Ohio, and was shipped in January 1991. The LPT stage 1 disk, SN MPOAP037 was installed in a new engine in April 1991. In February 2001, the first maintenance performed on the LPT stage 1 disk was when the disk had accumulated 16,254 hours time since new (TSN) and 4,645 cycles since new (CSN). At this time, the LPT rotor was disassembled to the piece part level and a Fluorescent Penetrant Inspection (FPI) of the LPT stage 1 disk revealed 5 cracks and damage to the seal serration. According to the maintenance records, the seal serration was blended to remove the cracks and damage and a post-repair FPI revealed no discrepancies. The LPT rotor was reassembled and installed in an engine in June 2001. In January 2002 the LPT stage 1 disk was removed after the engine was removed due to HPC blade damage. The LPT stage 1 disk had accumulated 17,306 TSN and 5,169 CSN. According to the maintenance records, the LPT rotor was not disassembled to the piece-part level. The inspection found no defects and no repairs performed. The LPT rotor was reinstalled in an engine in April 2002. In April 2004 the LPT stage 1 disk was removed when the LPT rotor was disassembled to separate the disks in order to replace the LPT forward shaft due to high time. The stage 1 disk had accumulated 20,882 TSN and 6,945 CSN. The LPT stage 1 disk was inspected (not de-bladed), which found uneven interlock wear on the outer shrouds of two blades. Those blades were removed and replaced and no other repairs were performed to the disk. The LPT rotor was reinstalled in engine SN 451-278, which was installed in position 3 on the accident airplane on July 12, 2004. At the time of the LPT disk separation, the LPT stage 1 disk had accumulated 23,970 hours TSN and 8,327 CSN. The LPT stage 1 disk is a life limited part with a life limit of 20,000 cycles. MANUFACTURING RECORDS OF LOW PRESSURE STAGE 1 DISK The disk forging was manufactured by Wyman-Gordan of Cleveland, Ohio, and was sent to GE for final machining. According to the “Dispatch Order” (the Dispatcher Order lists all operations involved in final machining, inspection and cleaning of the disk), finish machining of the disk forging commenced on June 7, 1988 at the GE facility in Evendale, Ohio. During the turning operation for the forward seal, the seal tooth diameter was over-machined, resulting in it being under the blue print minimum dimension, which was discovered during the bench inspection. This non-conformance was documented on a “deficiency record continuation sheet” on August 9, 1988. On May 3, 1989, the disk was released for dabber weld repair to restore the seal tooth diameter back to the original blue print dimension. According to the “non-conformance record continuation sheet” dated August 18, 1989, the subsequent inspection of that dabber weld repair revealed that the weld was not in the correct location and insufficient material build-up was applied, thus the disk was released for a second dabber seal tooth weld. The second dabber weld repair operation was completed on May 17, 1990. However, according to the “non-conformance record continuation sheet” and the oven bake records, distortion/warping of the disk was observed during the subsequent post weld stress relief process and the heat treat process was eventually aborted. Inferred from the “non-conformance record continuation sheet”, the disk was improperly insulated resulting in the observed warpage but was corrected and according to the oven bake records dated March 26, 1990, another attempt to stress relieve the disk was made. The oven bake records indicate that three thermocouples recorded a temperature 11ºF over the maximum allowable temperature, and according to the “non-conformance record continuation sheet,” the disposition was that the part should be continued to be processed as-is. The turning operation for the repaired seal tooth occurred on April 17, 1990, and the diameter was measured to be within the blue print specification. However, the follow-on audit inspection as recorded on the “deficiency record continuation sheet”, indicated that the seal tooth had five dents in a 1-inch area along the seal outer diameter and that the weld had not been cleaned up on the aft side of the seal. Hand-written notes on the “deficiency record continuation sheet” show that the dents were straightened and polished and signed off as acceptable on May 10, 1990. On January 24, 1991, the disk final inspection was performed, and the disk was subsequently shipped for use. LPT STAGE 1 DISK DABBER WELD REPAIR PROCEDURES The dabber weld repair of seal serrations is permissible for CF6-6, -50, -80A LPT stage 1 disks. The repair was introduced into the CF6 Engine Manual in August 1987. The CF6-6 Engine Manual dabber weld procedure permits the serration tip to be machined down to the designated “In-Process” dimension in order to clean up any cracks or damage to provide a uniform surface wide enough to initiate the weld buildup. A caution is included to prevent over machining the seal tooth tip and it states, “CAUTION: If part does not completely clean-up to within in-process dimensions, the part is not repairable per this repair.” The repair does not authorize the removal of the entire serration and weld build-up to the specified height; however, from the metallurgical results, the weld features of the event disk showed that the forward arm region was welded contrary to the MRB procedures and the CF6-6 EM procedures. CF6 OTHER DABBER WELD REPAIRED PARTS GE performed a search of CF6 LPT stage 1 disk manufacturing records for disks that may have been dabber welded at manufacture. The result of the search identified four CF6-6, one CF6-50, and no CF6-80A model LPT stage 1 disks having had their seal tooth dabber welded at manufacturing. All of the CF6-6 disks were repaired during the same time period, from 1988 to 1991, while the CF6-50 disk was repaired in 1993. Of the four CF6-6 disks, two were scrapped prior to shipping while the other two were shipped and entered into service, one of which was the event disk. The other CF6-6 disk was removed from service in January 1993 during a shop visit due to multiple FPI indications and was shipped to GE for evaluation. Micrographs of the disk revealed that the weld had extended the full length of the tooth and into the forward face of the arm, similar to the event disk. Review of the maintenance records revealed that, similar to the event disk, it had undergone two dabber weld repairs to restore the seal tooth diameter. Examination of the detailed routers for both the event disk and this other disk revealed that they were both welded at almost the same time. Both disks were first weld repaired within five weeks of each other while their second weld repair was performed within the same week. The CF6-50 disk that had a seal tooth serration repair was removed from service in July 2006 and was sent to GE for evaluation. Micrographs of the disk revealed that the weld did not extend into the arm but was found in the seal tooth area only, with the weld orientation in the radial direction only. Review of the maintenance records revealed that the disk was weld repaired once, with no reference to multiple repairs or weld passes. As the airplane climbed through about FL260, the No. 3 engine low pressure turbine (LPT) stage 1 disk burst when a fatigue crack that initiated in the forward seal tooth serration arm ruptured. Fragments from the burst disk partially severed the LPT case and struck the fuselage in multiple locations. The airframe vibration increased to a severe level until the LPT rotor separated from the engine, about 14 minutes 30 seconds after the uncontained LPT stage 1 disk burst, when the high pressure turbine aft shaft and the fan mid shaft failed due to contact damage and torsional loading. Metallurgical examination of the portion of the LPT stage 1 disk that was imbedded in the right wing found that the disk burst due to a mixture of high cycle and low cycle fatigue. The fracture was located in the vicinity of the forward seal tooth serration, although the exact fracture origin could not be identified because the entire seal tooth and part of the seal tooth arm were missing at the fracture location. Examination of the forward seal tooth serration revealed the presence of two previous weld repairs that extended into the forward arm. Review of the manufacturing records revealed that the disk had been dabber weld repaired twice to correct anomalies and non-conformances found during the original manufacturing process. The dabber weld procedures did not authorize welding in the forward seal tooth serration, therefore, the weld repair performed on the failed LPT stage 1 disk did not conform to the approved repair instructions. Based on testing and analysis of sample, post-manufacture dabber weld repaired disks, the suspect population of CF6 disks was confined to those that had been repaired at original manufacture. Records indicated that, in addition to the accident disk, four other disks had been dabber weld repaired at original manufacture to correct forward seal tooth serration anomalies and non-conformances. None of these disks are any longer in service. Based on the estimated crack propagation rate, it is likely that either no crack existed at the last inspection of the disk or the crack size was below the reasonably detectable threshold. After the disk burst, the flight crew completed all required checklists; however, the crew could not manually shutdown the affected engine because the burst disk fragments had severed the fire shutoff valve cable, the power off cable, and the emergency shutoff cable. The engine eventually shutdown on its own due to airflow disruptions within the engine. The flight crew’s procedural coordination and use of available resources, consistent with cockpit resource management best practices, helped to ensure the safe landing of the airplane. Source: NTSB Aviation Accident Database (Pre-2008 Archive) Retrieved: 2026-02-12
Verbatim from NTSB's published report. Source file
NTSB_2006_DCA06FA033.txt.
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