Experimental and Numerical Crack Initiation Analysis of the Compressor Blades Working in Resonance Conditions

This paper presents the results of a complex experimental and numerical crack initiation analysis of the helicopter turbo-engine compressor blades subjected to vibrations. A nonlinear finite element method was utilized to determine the stress state of the blade during the first mode of transverse vibration. In this analysis, the numerical models without defects as well as those with V-notches were defined. The quality of the numerical solution was checked by the convergence analysis. The obtained results were next used as an input data into crack initiation (ε-N) analyses performed for the load time history equivalent to one cycle of the transverse vibration. In the fatigue analysis, the different methods such as: Neuber elastic-plastic strain correction, linear damage summation and Palmgreen-Miner rule were utilized. As a result of ε-N analysis, the number of load cycles to the first fatigue crack appearing in the compressor blades was obtained. Moreover, the influence of the blade vibration amplitude on the number of cycles to the crack initiation was analyzed. Values of the fatigue properties of the blade material were calculated using the Baumel-Seeger and Muralidharan methods. The influence of both the notch radius and values of the UTS of the blade material on the fatigue behavior of the structure was also considered. In the last part of the work, the finite element results were compared with the results of experimental vibration HCF tests performed for the compressor blades.