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Compressor and their blades are one of the most important parts of gas turbines. Based on recent reports, failure of compressor’s blades was one of the major cause in malfunctioned t56 gas turbines. In this study, propagation rate of a crack within the compressor blade of a T56 jet engine has been investigated. To this end, centrifugal and aerodynamic forces acted upon the blade has been calculated and their corresponding stress field has been simulated in ANSYS software. Spots at the maximum risk of foreign object damage and corrosion had been located, and their bending and tension stresses had been calculated via employed simulation. Subsequently, an initial half elliptical crack has been created on all of previously located spots, and their stress intensity factor using Raju-Newman method has been determined. Finally, by using Paris law fatigue life and crack growth rate of each cracks has been extracted, individually. Results indicate a drastic decrease in fatigue life of blades when crack located close to the blade’s root. Furthermore, cracks located on the suction surface has remarkably shorter fatigue life than those which are located on pressure surface, in comparison.