Flight Line Mitsubishi MU-2B-36 incident

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.

Verbatim NTSB analysis and probable cause from the NTSB dataset