Abstract: Shock oscillations are common flow phenomena encountered in the unstarting processes of hypersonic inlets. They can significantly reduce the airflow capturing and compression efficiency and generate severe unsteady aerodynamic loads. These are highly detrimental to the fight safety of hypersonic vehicles. Aiming at the control of shock oscillation flows, the effects of vortex generators on the oscillatory flows of an axisymmetric hypersonic inlet are studied experimentally. Both started flows and oscillatory flows are investigated with synchronized high speed schlieren imaging and transient surface pressure measurement, as no vortex generators, 0.5 mm and 1 mm thick vortex generators fixed on the inlet forebody. According to the experimental results, minor effects are exerted on the main flowfield and wall pressure of the started flows for vortex generators within 1 mm thickness. However, the vortex generators can substantially reduce the scale of external separations, shorten the oscillatory period and increase the time- averaged pressure magnitude of the oscillatory flows. The effects of vortex generators enhance with the increase of their thickness, and the oscillatory period is shortened from 4 ms for no vortex generator cases to 3.13 ms for 1 mm thick vortex generator cases. In addition, the 0.5 mm thick vortex generators can generally reduce the oscillatory amplitude of surface pressure of the inlet duct, of which decreasing percentage can reach 23%. According to the analysis of the schlieren images and the surface pressure signals, effects of the vortex generators are exerted through streamwise vortexes in the wake flows, including the disturbances of streamwise vortexes exerted on the downstream boundary layers and the interactions between streamwise vortexes and separation regions.