Two damaged I stage compressor rotor blades along with two locking pins belonging to an aeroengine were forwarded to the laboratory for investigation. Fractographic examination showed that the blade number 24 had fractured by sudden overload such as impact. The fracturing of this blade was secondary in nature and subsequent to the primary failure. Evidences suggest that the dislodgement of blade number 5 was responsible for the engine failure. The dislodgement of the whole blade without any breakage is not possible unless the locking mechanism fails. It appears most probable that in this particular case, this was possible due to an assembly error wherein a locking pin of length 27.0 mm was used against the specification of 32.0 mm.