A study on the transitional boundary layer with arbitrary pressure gradient under various upstream conditions is very important for engineering applications like the performance predictions of the turbomachineries under various and strong disturbances. Experimental data on the transitional boundary layer for real cascades of the turbomachinery are rare because of difficulties in boundary layer experiments. Flow on NACA0012 airfoil is more similar to the real case than that on the flat plate with which many researches are done. The data of the transitional flow on the airfoil could be used to verify or to develop a turbulence model for numerical simulations. The experiment was performed with two cases of Reynolds number at a=0

and one case of Reynolds number at a=5

. The measured data are the transition length and the wall shear stresses. These two characteristic values are measured within 25%~90% of the airfoil chord by Computation Preston tube Method(CPM) proposed by Nitsche et al.(1983). At a=0

, transition occured at 70% and 55% of chord length when R

=6*10

and 8* 10

, respectively. It started when R {\theta}=500 regardless of R

, and ended when R {\theta}=750, and 850 respectively. The transition length was 15~20% of the chord length. At a=5

(R

=6*10

), boundary layer on the pressure side does not undergo transition, but on the suction side transition occured at .chi.

/c=0.16 and ended at .chi.

/c=0.22.c//c=0.22./c=0.22.c//c=0.22.