Gas turbines are used in trains, ships, tanks, and alongside steam turbines and in power stations to generate power. During every start-up and shutdown of an aviation gas turbine, the compressor blades are subjected to centrifugal, gas bending and vibratory loads. This repeated loading and unloading can reduce the life of compressor blades. With blading problems accounting for as many as 42 percents of the failures in gas turbines. Previous works on compressor blades have focused mainly on fatigue life estimation in the vicinity of foreign object damage. Geometric modeling of the blade was done using CAD tool CATIA. Static stress analysis was carried out to ascertain the critical region or crack zone of the blade. The maximum Von - Mises stress was found at the fillet region near the root of the blade The results of Finite element analysis showed that maximum von-mises stress was found at the 12 o‟ clock fillet region due to an influence of centrifugal stresses. The maximum Mode I stress intensity factor of 61.4 Mpa√m was found at the surface interception point, for a crack length of 6mm and crack depth of 2.4mm. With the increase in rotational velocity of 5000 rpm, 10000 rpm, and 20000 rpm, Fatigue crack length growth rate was estimated to be a 1.95 x 10-09, 1.65 x 10-07 m/cycle and 4.15 x 10-05 m/cycle and the fatigue crack propagation life was estimated to be of 4.3 x 108 cycles, 7.3 x 106 cycles and 5 x 103 cycles respectively.