Chicago, IL, USA
N950MA
MITSUBISHI MU-2B-36
The airplane performed an uneventful single-engine landing after experiencing an uncontained No. 2 (right) engine failure during final approach. The pilot reported feeling a shudder, hearing a loud bang, and observing a fire ball coming from the right engine during landing gear extension. Postflight inspection revealed an exit hole in the inboard side of the right engine cowling. There was a similar sized penetration in the right wing above and inboard of the cowling hole. A rupture was found on the left side of the engine plenum case, in line with the airframe damage. The appearance of the case and the corresponding cowling and wing penetrations indicated that a single, high-energy rotor fragment had exited the engine case in an upward and slightly forward trajectory. Although the 2nd stage turbine wheel was not recovered, the physical evidence, including the axial location of the rupture, the intact condition of the 1st stage turbine wheel, and the absence of multiple case penetrations, suggested that a 4 to 5-inch piece of rim material separated from the 2nd stage turbine wheel and exited the side of the engine. According to Honeywell, a rim fragment separating from the 2nd stage turbine wheel will typically lack the energy for a direct radial release, and the engine structure showed marks and damage consistent with a released 2nd stage rim fragment striking the 3rd stage turbine stator assembly forward rib support casting and deflecting forward to exit at the rupture location between the 1st and 2nd stage turbine planes of rotation. The heavy curvic teeth smearing damage indicated gapping at the rotor curvic connections during operation, a severe condition that will result from gross rotor mass unbalance, such as the unbalance that results from the loss of a section of turbine wheel rim material. The shaft separation in plane with the 2nd stage turbine wheel curvics is also consistent with a rotor mass unbalance at the 2nd stage turbine wheel. The extensive non-uniform rotational damage is consistent with the loss of centering and the loss of operating clearances resulting from severe rotor unbalance and sudden shaft separation during operation. Honeywell’s materials analysis suggested that a section of wheel rim material separated as the result of low cycle fatigue (LCF) cracks originating at the 2nd stage turbine wheel stress-reduction "rivet" holes,and that fatigue cracks progressed to failure from long term exposure to excessive temperatures. Rivet hole fatigue cracks as a rim separation failure mode was not confirmed, as only a small amount of 2nd stage turbine wheel material was recovered, and examination of these fragments found no evidence of LCF. Although rivet hole fatigue cracks do not normally progress to failure withing wheel life limits, elevated operating temperatures will contribute to crack growth. Moderate to severe thermal damage was observed in the engine's other turbine components with similar time in service, and the metallurgical evidence supported a finding that 2nd stage turbine vane airfoils were exposed short term to temperatures exceeding 2000 degrees F. No material defect or anomaly was found, however, due to missing critical material, material defects, anomalies or other failure mechanisms could not be ruled out. The engine failure most likely resulted from a partial rim separation of the 2nd stage turbine disk during operation. The rim separation failure mechanism, or potential contributing factors, could not be identified because critical parts, including the 2nd stage turbine disk, were not recovered, despite several searches.
HISTORY OF FLIGHT On January 12, 2007, about 0047 Central Standard Time, a Mitsubishi MU-2B-36 airplane, N950MA,powered by two Honeywell TPE331-6-252 engines, made an uneventful single-engine landing after experiencing an uncontained No. 2 (right) engine failure while on approach to Midway Airport (MDW),Chicago, Illinois. The pilot reported feeling a shudder, hearing a loud bang, and observing a fire ball coming from the right engine during landing gear extension. No injuries to the pilot or persons on the ground were reported. The airplane sustained minor damage. The flight was operated by Flight Line, Inc. doing business as American Check Transport, under the provisions of 14 Code of Federal Regulations Part 135, as a cargo flight from Milwaukee, Wisconsin to Chicago, Illinois.Visual flight rules conditions prevailed at the time of the incident. AIRCRAFT INFORMATION A review of the engine’s maintenance records showed that new 1st, 2nd, and 3rd stage turbine wheels and a new 2nd stage turbine stator assembly were installed in the engine 2,841 hours and 2,646 cycles before the incident. The records place the turbine wheels at approximately midlife relative to the wheel life limits. WRECKAGE AND IMPACT INFORMATION A postflight inspection found an exit hole in the inboard right engine cowling and a similar sized penetration in the right wing above and inboard of the cowling hole. Inspection through the engine exhaust found that most of the turbine components downstream of the 1st stage stator assembly were missing. The remaining structure was battered, with extensive circumferential scoring. A ground search was conducted along the final approach path to MDW, and engine parts including the 1st stage turbine wheel, center curvic coupling with sections of the main and torsion shafts, and a section of the 2nd stage turbine stator ring, were recovered. Most of the parts were found on rooftops or on the ground; however, the 1st stage turbine wheel penetrated the roof of a one-story home, where it was found by the homeowner on the floor of a bedroom. A second ground search was conducted and several additional parts were located, but the 2nd and 3rd stage turbine wheels were not recovered. The engine was examined at a Honeywell facility in Phoenix, Arizona. There was an approximate 5.5 inch (circumferential) by 1.8 inch (axial) hole in the left side of the plenum case just forward of the 2nd stage turbine wheel plane of rotation. There were no additional case penetrations. The edges of the rupture were curled outward with the most displacement at the top and forward edges.The outer transition liner, outer combustor skirt, and inner combustion skirt were torn inboard of the plenum case rupture and the material at the damage sites was torn and curled outward. The outer combustor skirt was also battered and exhibited circumferential scoring along the 3rd stage turbine plane of rotation. The 1st stage compressor impeller curvic teeth were severely smeared. The radial shroud contour of the 2nd stage compressor housing and shroud was deformed radially outward and scored. The 2nd stage compressor impeller curvic teeth were heavily damaged. The 1st stage stator assembly vane airfoils showed extensive trailing edge (TE) cracking, missing material, and burn through. The 1st stage turbine wheel, which penetrated the house, was heavily damaged, including both the forward and aft curvics. Second stage stator sheet metal material remained attached at the 3rd stage stator assembly and a section of the sheet metal outer support remained assembled in the forward end of the rear bearing housing. 2nd stage stator fragments were recovered loose in the engine and along the airplane’s flight path. Most of the stator assembly inner wall casting was recovered, including a 200 degree arc with intact vanes. The vane airfoil TEs showed extensive thermal damage. There were heavy,non-uniform rubs along the forward edge of the outer ring and on both sides of the inner shroud inner diameter (ID). None of the 2nd stage stator assembly inner support or seal structure was recovered. The center curvic coupling was found with separated segments of the shouldered (main) and torsion shafts inside. The main shaft was cocked at the coupling internal mating splines, and these splines were severely deformed. There were heavy, non-uniform circumferential rubs on the coupling outer diameter and knife edge seal tips. The coupling forward and aft curvic teeth were scored and smeared. The 2nd stage turbine wheel was not recovered. 17 fragments, including 12 blade platforms recovered from inside the engine, were identified as 2nd stage blade material. The 3rd stage stator assembly showed extensive damage. The outer ring remained installed in the engine. There was a shiny arc of displaced material on the casting, just aft of the plenum rupture. The 3rd stage turbine wheel was not recovered. 14 blade platforms recovered from inside the engine were identified as 3rd stage blade material. The aft curvic coupling exhibited heavy scoring along its aft diameter. The (forward) curvic teeth were damaged but were not smeared. The rear bearing housing damage included extensive gouging and non-uniform circumferential scrapes across the remaining housing inner surfaces, including over the inner bearing support strut attachment fractures and along the 2nd stage turbine stator outer support remnants that remained attached to the housing forward end. The tailpipe showed aftward spiral patterns of contact marks in the direction of engine rotation consistent with parts exiting the engine through the exhaust. The main and torsion shafts, which are concentric, were separated in plane with the center curvic coupling forward face. The main shaft was also separated 2.5 inches aft of the center curvic coupling aft end, and the torsion shaft was separated 4.5 inches aft of the center curvic coupling aft end. The shaft sections aft of the coupling, which were roughly in line with the 2nd stage turbine wheel position, were bent and deformed. TESTS AND RESEARCH Parts of interest were submitted to a Honeywell materials lab in Phoenix, Arizona. Examination of 2nd stage turbine stator assembly fragments found low cycle fatigue (LCF) cracks at several vane TEs, and at the casting braze joints of the sheet metal outer and aft supports. LCF cracks were also detected in the rivet holes on the 1st stage turbine wheel; none of these cracks exceeded 0.110-inch in length. No fatigue indications were found on the fracture surfaces of the 2nd stage and 3rd stage wheel blade fragments recovered from inside the engine. All other fracture surfaces displayed overload features. The shaft fractures showed overload produced by a combination of bending and rotational motions. Complete gamma prime resolutioning with secondary diffusion of the aluminide coating was observed in the 2nd stage turbine stator vane airfoil inner shroud suction side, adjacent to the LE,consistent with short term exposure to metal temperatures exceeding 2000 degrees F. The 2nd stage turbine stator outer support also showed “fairly heavy grain boundary and matrix precipitates.” According to Honeywell,this structure indicated that the outer support had experienced a long time exposure to 1200 degree F to 1400 degree F temperatures. No material defects were found. The material compositions were as specified. The Honeywell findings were reviewed by the National Transportation Safety Board materials lab. TPE331 turbine wheels are integrally bladed, with slots between each blade at the rim. Each of the slots ends with a circular “rivet” hole. The slots and holes reduce thermal and centrifugal stresses at the heated outer rim and extend turbine life. LCF cracks tend to develop at the rivet holes but normally do not progress to failure within wheel life limits. However, elevated operating temperatures will contribute to crack growth. TPE331 engine rotating elements include integral curvic-type couplings that act as both centering and driving devices. A shouldered (main) shaft serves as a tiebolt that maintains the tension required to ensure full stack engagement during operation.
Complete loss of No. 2 engine power due to a partial rim failure of the 2nd stage turbine wheel, resulting in No. 2 engine shutdown and turbine debris exiting the side of the engine. The wheel failure mechanism could not be determined due to missing turbine components that were not recovered despite two ground search attempts.
Source: NTSB Aviation Accident Database
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