N526MD
MCDONNELL-DOUGLAS DC-10-30
The airplane experienced an uncontained No. 2 engine failure during climb out after takeoff, and an uneventful landing was accomplished. A post-incident airplane inspection found that the No. 2 engine low pressure turbine (LPT) case had separated circumferentially in plane with the LPT stage 3 (S3) turbine and that all components aft of the separation were liberated. The engine's LPT S3 disk front spacer arm was fractured circumferentially at the spacer to disk rim fillet. Metallurgical examination of the fracture surfaces revealed high cycle, high-amplitude fatigue (HAF) cracks initiating from sites around 90 percent of the forward spacer arm inner diameter. Inspection of the high pressure turbine (HPT) found that one stage 1 (S1) blade was missing approximately 85 percent of its airfoil material. The investigation found that the HPT blade damage resulted in synchronous vibration forces that interacted with the engine LPT rotor system through a common bearing support and excited a bladed-disk mode response in the LPT S3 disk. The resonant amplitude of the alternating stress experienced by the disk resulted in bending loads that exceeded the material endurance limit and caused high cycle/high amplitude fatigue cracks to initiate along the LPT S3 disk forward spacer arm rim diameter. Once initiated, the cracks propagated rapidly through the spacer arm thickness and the individual cracks joined to form a single circumferential crack, resulting in the separation of the disk at the forward spacer arm. The freed aft portion of the LPT rotor accelerated and penetrated the engine case, releasing high-energy debris as disk fragments and all of the engine components aft of the LPT S3 nozzles were liberated.
HISTORY OF FLIGHT On March 26, 2009, a McDonnell Douglas DC10-30F, N526MD, experienced an uncontained No. 2 engine failure during climb out after takeoff from Eduardo Gomes International Airport, Manaus, Brazil (MAO). According to the flight crew, about 30 minutes after takeoff, at about 8,000 feet above ground level, the airplane's No. 2 engine oil pressure steadily decreased. The Captain shut down the engine and continued the flight. The flight was later diverted to José Maria Cordova International Airport, Medellin, Colombia, where an uneventful landing was accomplished. A post-incident airplane inspection found damage to the, tail strut, elevators, rudder, horizontal stabilizer, and the No. 2 engine pylon and nacelle. The aft end of the No. 2 engine had separated roughly in plane with the low pressure turbine (LPT) case had separated and all components aft of the separation were missing. The flight crew did not report any unusual engine vibration preceding the event. The airplane was operated by Arrow Cargo under the provisions of 14 Code of Federal Regulations Part 21 as a regularly scheduled cargo flight from MAO to El Dorado International Airport, Bogota, Columbia. There were no injuries. Liberated engine parts impacted a densely populated area near Manaus, and 22 homes were reported damaged. DAMAGE TO AIRPLANE There was no major structural damage to the airplane. There was extensive tearing damage to the No. 2 aft pylon fairing, aft engine support mounting bracket, and the No. 2 engine core cowl doors. There were perforations and dents in the rudder, the upper and lower skins of the left and right wing elevators, and the right horizontal stabilizer. AIRCRAFT INFORMATION The airplane is a McDonnell Douglas DC10-30F, powered by 3 General Electric (GE) CF6-50E2 turbofan engines. The airplane was registered to Miami Leasing, Inc. and operated by Arrow Air Holdings Corporation doing business as Arrow Cargo. The incident engine, which was installed in the airplane's No. 2 (tail) position, was owned by AAR Corporation and leased to Arrow Cargo. Arrow Cargo maintenance records showed that the engine had accumulated 83,895 hours since new and 21,904 cycles since new and had been overhauled 7,883 hours and 2,703 cycles before the incident. The records also showed that No. 2 engine high pressure turbine (HPT) borescope inspections (BSIs) performed on May 5, 2008, August 8, 2008, and October 28, 2008, had found that the engine HPT was in airworthy condition per airplane maintenance manual (AMM) limits. FLIGHT RECORDERS The airplane was equipped with a Honeywell Model 980-4100 flight data recorder, which records 22 flight parameters. The data showed that the No. 2 engine fan speed began to decrease as the throttle was increased to climb power and continued to decrease for about 65 seconds, reaching 14 percent RPM and remaining below 27 percent until the FDR data ended. ENGINE INVESTIGATION The Brazilian Air Force Centro de Investigação e Prevenção de Acidentes Aeronáuticos (CENIPA) recovered many of the liberated parts. The United States accepted delegation of the investigation from CENIPA as the State of Registration, Design, and Manufacture of the engine and airplane, and the National Transportation Safety Board (NTSB) arranged to have the engine and the recovered parts shipped to the United States for further investigation. The engine was examined at Kelly Engine Center, San Antonio, Texas. An external inspection found no obvious engine damage forward of the turbine midframe (TMF); aft of the TMF there was localized crush damage and the LPT stator case was separated 360 degrees roughly in plane with the LPT S3 disk plane of rotation. The LPT S3 disk rim, web, bore, and aft spacer arm; 39 LPT S3 nozzle segments; all of the LPT stage 4 (S4) nozzle segments; the LPT S4 disk; and the turbine rear frame had liberated. The LPT S3 disk forward spacer arm was attached to the LPT forward shaft and was fractured circumferentially approximately 1.54 inches aft of the flange forward face, which is the approximate location of the fillet between the spacer arm and the disk rim. The engine was disassembled and a rotor check balance found that the total HPT rotor unbalance was 1,213 gram-inches; the CF6-50 Engine Manual HPT rotor unbalance limit is 40 gram-inches. A single HPT S1 blade was severely eroded, with approximately 85% of its airfoil material missing. The combustor positioning pins and bushings were severely worn, and combustor and HPT S1 nozzle components showed localized wear consistent with aftward shifting of the combustor during operation. The exhaust gas temperature (EGT) system was in poor condition, with the outer sheaths of several probes partially or completely missing. Metallurgical examination of the LPT S3 disk fracture surfaces revealed features consistent with a high-amplitude per-revolution stimulus that resulted in high cycle, high-amplitude fatigue (HAF) crack propagation over approximately 90 percent of the fracture. The remaining 10 percent of the fracture showed overstress features. The cracks propagated from initiation sites spaced about 0.1 inch to 0.2 inch apart around the inner circumference of the spacer arm and joined to form a single circumferential crack, leading to disk separation. The disk material met all specifications with the exception of a small deviation in the as-large-as (ALA) grain size requirement, which did not play a role in the fatigue fracture. RESEARCH AND TESTS CF6-50 HPT rotor unbalance-induced LPT S3 bladed-disk mode According to GE, when the CF6-50 experiences a high level of HPT rotor unbalance, the resulting synchronous vibration forces can interact with the engine’s LPT rotor through a common bearing sump and excite a bladed-disk mode response in the LPT S3 disk. The resonant frequency experienced by the LPT S3 disk in this mode will result in forward spacer arm bending loads that can exceed the fatigue limit of the material and result in HAF crack initiation. The CF6 LPT S3 disk resonance response to HP rotor unbalance was first identified in the GE CF6-6 engine, which shares the CF6-45/-50 type certificate. The CF6-6 experienced four uncontained LPT S3 disk forward spacer arm separations between 1975 and 1978 due to HP rotor unbalance. As a result, GE redesigned the CF6-6 LPT S3 disk so that an HP rotor unbalance condition would not excite the LPT S3 disk and result in disk failure. The CF6-50 engine experienced 10 instances of LPT S3 disk forward spacer arm cracking between 1973 and 2009. Eight of the cracked CF6-50 disk forward spacer arms were discovered during shop-level inspections when LPTs were disassembled for unrelated reasons, such as engine model conversion or the replacement of life-limited parts. In the remaining two cases, disk cracks progressed to failure, leading to in-service uncontained engine failures. Field detection of HPT rotor unbalance The DC10-30 airplane was certificated with an engine vibration monitoring (EVM) system. EVM systems are safety systems that measure engine HPT rotor unbalance, in order to alert flight crews of impending engine failure. However, the DC10-30 EVM system was considered marginal in its ability to detect CF6-50 HPT rotor unbalance, and, when the CF6-45/-50 engine was being certified, the FAA placed a special condition (No. 33-36-EA-9, dated November 8, 1971) on the CF6-45/-50 type certificate that required GE to show that “the engine would operate without inducing detrimental stresses in any engine part while operating with an increased vibration level, such as that which might result from one or more broken or missing rotor blades, if the increased vibration level cannot be detected in flight.” The FAA considered this special condition satisfied with its acceptance of a GE certification report that concluded that CF6-45/-50 engine unbalance levels high enough to have detrimental effects would be easily detected by perceived noise and vibrations in the cabin area and in the controls, enabling the flight crew to take corrective action. EVM systems were decertified in many CF6-50-powered airplanes as permitted by Douglas Aircraft Company All Operator Letter 10-1692, “Engine Vibration Monitor (EVM) Deletion,” dated February 21, 1983. HPT rotor blades are expected to deteriorate (lose blade airfoil material) between overhaul intervals due to the operating stresses they experience. Deterioration is detected by periodic borescope inspections for blade material loss, and by the monitoring of EGT trend data. The blade material loss is usually symmetrical and does not significantly affect rotor balance. However, when the material loss is localized, detrimental rotor unbalance will occur that can result in synchronous, per-revolution vibration. According to GE, the loss of 85% of one HPT blade airfoil will result in a level of HPT rotor unbalance sufficient to excite a bladed-disk mode in the LPT S3 disk. Arrow engine performance trend data A review of Arrow Cargo’s exhaust gas temperature trend data for the incident engine found that the engine parameter data lacked fidelity to the extent that no identifiable performance trend could be determined. ADDITIONAL INFORMATION The GE CF6-50 LPT S3 disk has a cycle limit of 12,400. At the time of the incident, DC10-30 AMMs required borescope inspection (BSI) of CF6-45/-50 HPT rotor blades every 450 flight cycles as part of the engine hot section BSI. As a result of this investigation, and the investigation of a similar event that occurred on July 4, 2008, the NTSB made the following recommendations to the Federal Aviation Administration(FAA) on May 27, 2010: 1. Immediately require operators of CF6-45/-50-powered airplanes to perform high pressure turbine rotor blade borescope inspections every 15 flight cycles until the low pressure turbine stage 3 disk is replaced with a redesigned disk that can withstand the unbalance vibration forces from the high pressure rotor. (A-10-98) (Urgent) 2. Require operators of CF6-45/-50-powered airplanes to perform fluorescent penetrant inspections of CF6-45/-50 low pressure turbine stage 3 disks at every engine shop visit until the low pressure turbine stage 3 disk is replaced with a redesigned disk that can withstand the unbalance vibration forces from the high pressure rotor. (A-10-99) 3. Immediately require General Electric Company to redesign the CF6-45/-50 low pressure turbine stage 3 disk so that it will not fail when exposed to high pressure rotor unbalance forces. (A-10-100) (Urgent) 4. Once General Electric Company has redesigned the CF6-45/-50 low pressure turbine stage 3 disk in accordance with Safety Recommendation A-10-100, require all operators of CF6-45/-50-powered airplanes to install the newly designed LPT S3 at the next maintenance opportunity. (A-10-101) On March 17, 2010, the FAA published immediate release Airworthiness Directive (AD) 2010-06-15, which required fluorescent penetrant inspection (FPI) of the CF6-45/-50 low pressure turbine (LPT) stage 3 (S3) disk under certain conditions and removal of the disk from service before further flight if found cracked, and initial and repetitive borescope inspections (BSI)of the high-pressure turbine (HPT) rotor stage 1 (S1) and stage 2 (S2) blades for wear and damage, including excessive airfoil material loss. On June 9, 2010, The FAA published AD 2010-12-10, which required initial and repetitive BSI of the HPT rotor S1 and S2 blades for wear and damage, including excessive airfoil material loss, FPI of the LPT rotor S3 disk under certain conditions and removal of the disk from service before further flight if found cracked, and repetitive exhaust gas temperature (EGT) system checks (inspections). On February 7, 2011, the FAA published AD 2011-02-07, which superseded AD 2010-12-10 and required on-wing inspections of the HPT rotor S1 and S2 blades, exhaust gas temperature (EGT) system inspections, engine core (HPT)vibration surveys, and an ultrasonic inspection (UI) of the LPT rotor S3 disk forward spacer arm. That AD also required fluorescent-penetrant inspection (FPI) of the LPT rotor S3 disk under certain conditions and removal of the disk from service before further flight if found cracked. In June 2011, the FAA certificated a new-design CF6-50 LPT S3 disk that has improved tolerance to HPT unbalance forces. On January 31, 2012, the FAA published AD 2012-02-07, superseding AD 11-02-07 and AD 2011-18-01, which established a new lower life limit for the old-design LPT rotor stage 3 disk and introduced a draw down plan for the removal of the old-design disks from service. The AD also retained the requirements of the superseded ADs and adds an optional LPT S3 disk removal after a failed HPT blade BSI or a failed engine core vibration survey.
Failure of the low pressure turbine stage 3 disk due to a design that is vulnerable to high pressure turbine unbalance-induced synchronous vibration that cannot be detected in flight, and the subsequent uncontained engine failure.
Source: NTSB Aviation Accident Database
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