An experimental study was conducted in order to investigate unsteady boundary layers for a pitching airfoil. An NACA0012 airfoil sinusoid-pitched at quarter chord was employed, and its mean angle-of-attack and oscillation amplitude were $0^{circ}$ and $6^{circ}$, respectively. To explore the unsteady boundary layers, smoke-wire visualization and surface-mounted probe measurements were pursued for three different cases, varying with Reynolds numbers ($Re_c=2.3{ imes}10^4,;3.3{ imes}10^4$, and $4.8{ imes}10^4$). A reduced frequency of 0.1 was identically set in all cases. Results show that in the presented Reynolds number range, the separation bubble dependent on both angle-of-attack and Reynolds number was observed, accompanied with unsteady laminar separation after reattachment. The unsteady laminar separation occurred at the saddle point, which was formed by the two vortices, the wall, and the external flow, and it was independent of reverse flow. This result indicates that the unsteady laminar separation occurs during the process of transition after the reattachment of separated boundary layer for an unsteady flow. The reverse flow observed over the trailing edge significantly interacted with the trailing edge vortex that rotates in the streamwise direction. This trailing edge vortex prevents the uppermost of the reverse flow from reaching to the unsteady laminar separation point during the upstroke, and this induces that the boundary layer breakdown does not occur in spite of the occurrence of laminar separation. The discrete vortices are formed by unsteady laminar separation, and its formation is ultimately affected by the Reynolds number. Consequently, it is obvious that the unsteady boundary layers are ultimately sensitive to Reynolds number in a low Reynolds number regime.