Event Summary
During final approach, the aircraft experienced an uncontained failure of the No. 2 (right) engine while extending the landing gear. The pilot reported feeling a shudder, hearing a loud bang, and observing a fireball emanating from the right engine. The aircraft performed an uneventful single-engine landing.
Damage Findings
Postflight inspection revealed an exit hole on the inboard side of the right engine cowling. A similar-sized penetration was found in the right wing, above and inboard of the cowling hole. A rupture was discovered on the left side of the engine plenum case, aligned with the airframe damage. The appearance indicated that a single, high-energy rotor fragment had exited the engine case in an upward and slightly forward trajectory.
Further examination showed heavy curvic teeth smearing damage, indicating gapping at the rotor curvic connections during operation. The shaft separation occurred in plane with the 2nd stage turbine wheel curvics. Extensive non-uniform rotational damage was consistent with loss of centering and operating clearances due to severe rotor unbalance and sudden shaft separation.
Metallurgical Analysis
The 2nd stage turbine wheel was not recovered. 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 typically lacks energy for a direct radial release; engine structure marks and damage were consistent with a released fragment striking the 3rd stage turbine stator assembly forward rib support casting and deflecting forward to exit at the rupture location.
Honeywell's materials analysis suggested that a section of wheel rim material may have separated due to low cycle fatigue (LCF) cracks originating at 2nd stage turbine wheel stress-reduction rivet holes. Fatigue cracks may have progressed from long-term exposure to excessive temperatures. However, this failure mode was not confirmed, as only a small amount of 2nd stage turbine wheel material was recovered, and examination found no evidence of LCF. Elevated operating temperatures can contribute to crack growth. Moderate to severe thermal damage was observed in other turbine components with similar service time. Metallurgical evidence supported that 2nd stage turbine vane airfoils were exposed short-term to temperatures exceeding 2000 degrees F. No material defect or anomaly was found, but due to missing critical material, defects or other failure mechanisms could not be ruled out.
Probable Cause
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.