The role played by short-haul aircrafts is critical in air transport network. With service missions of approximately 3 hours or distances not exceeding 2000 km, the compressor turbine (CT) blades of their engines experience intense cyclic thermo- mechanical stresses from short but high flight turnovers, changes in power settings, starts and stops, in adherence to the engine’s preset operation limits, overexploitation and not sticking to preset flight environments.Thermo-mechanical stresses implacably account for the greatest degradation of CT blades. From continual exposure to this severe operational environment, the CT blades often catastrophically fail without warning. This study, therefore, investigates thermo-mechanical damage evolution a typical high pressure (HP), PT6A-114A engine CT blade. An assimilative investigating approach was adopted. The CT blade was modeled for the evolution of thermo-mechanical stresses in an environment that mimics the operational conditions using commercial ANSYS tools in the transient regime. A detailed microstructural and metallographic characterization was then performed using energy dispersive spectroscopy-scanning electron microscopy (EDS-SEM) on the substrate material. Finally, mechanical testing for residual micro hardness and equivalent yield strength executed. The modeling results revealed that the CT blade experienced greater deterioration at the tip and lesser on the airfoil and least on the base. In agreement with the modeling results, the EDS-SEM results established that the substrate material had degraded from the effect of creep and fatigue with greatest damage on the tips, lesser on the airfoils and least at the bases. The experimental microhardness values for the transverse sections had fallen by 28.5%, 23% and 20.5% at the tips, airfoil, and base respectively and 25.1%, 22.6% and 20% for longitudinal section relative to the typical hardness values of HV 440 for Ni-base superalloy CT blades before exposure to service.

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