An experimental investigation on forced response of a first stage compressor mistuned bladed disc rotor is performed in an operational aero-engine. It is desired to evaluate the maximum blade amplitude response of the bladed disc for the existing blade mistuning pattern depending on blade mode, exciting engine order, and aerodynamic influences. Finite element analysis on a single cantilever blade is performed and the possible resonances are estimated. Bench tests on a stand-alone blade are conducted to verify the finite element predictions. Experiments are conducted on a full-scale core engine using blade tip timing system. Blade tip timing system utilizes noncontacting sensors installed on the engine casing which provides vibration responses of all the blades in a rotor. The vibration responses obtained from blade tip timing system is validated using simultaneously measured strains from a few blade-mounted strain gages. A good correlation is observed between the predicted and measured blade responses. The effect of blade mistuning on the amplitude resonant response of the bladed disc is explored experimentally. It is observed that the maximum amplitude response of the blades due to forced resonant excitation is far below the fatigue limits for the prevailing blade mistuning pattern. The present exercise of quantifying the blade resonant responses through efficient implementation of the blade tip timing system is crucial to understand and thus avoid high-cycle fatigue, which is one of the most serious problems in aero-engine development. © IMechE 2015.