Based on the analysis, the failure of the 1st stage compressor blade from the subject GT occurred by a high-cycle fatigue (HCF) mechanism. The root cause of the failure was attributed to an internal metallurgical anomaly near the airfoil leading edge. Fatigue cracks initiated from the anomaly region and propagated towards airfoil mid-chord until final tensile overload separation occurred.
Therefore, manufacturing process of the GT blades should be well monitored and controlled to avoid residual stresses or surface defects. Random checks on blades from each forging heat lot will reduce such risks. Operator should implement a rigorous on-condition monitoring of the GT rotor and spot any side bands close to the BPF excitation peaks. Load drop or transient regime of the turbines increases vibration amplitude when harmonic rotor frequencies are interacted with blade’s natural frequencies. Thus, the number of engine’s start-up and shutdown of the GT shall be reduced to the minimum. Rou- tine boroscopic inspection of the GT rotor should focus on the stress concentration area of the blade located near the con- nection region of the airfoil and the root as shown by the FE analysis.
The static loading analysis performed in this work needs to be completed by dynamic analysis to assess the blade resis- tance to fatigue induced by repeated/fluctuated loads and the aerodynamic cyclic stresses. A thermo-elastoplastic behavior of the blade material is required for identifying the blade’s strain-life fatigue crack initiation and propagation.