One way to improve the thermal efficiency of gas turbine is increasing the turbine inlet temperature. Various types of cooling technology have been applied in gas turbine blade either external or internal cooling. Based on the manufacturing process, the internal cooling is the easiest one. Modification of the number and size of the internal cooling passage plays important rule in blade design. It affects the amount of heat transfer, the temperature distribution and the gradient inside the blade and also the mass flow rate of the cooling air and thus resulting the overall efficiency of the turbine. Therefore objective of the present study is modifying number and the size of holes cooling channel of NASA C3X Turbine blade by using ANSYS 19.2. The variation of mass flow rate of ṁ -10%, ṁ + 10%, and ṁ + 20% from the normal design of C3X turbine blade have been applied. The results show that the modified cooling holes has an adverse effect than the normal turbine blade. The modified cooling holes increases the blade surface temperature by 3.29% which is equivalent to 20.85 K, reduces the heat transfer coefficient by 16.43% which is equivalent to 72.85 W/m2K and reduces the total heat transfer rate by 13.51% which is equivalent to 648.14 W. But in another side by increasing the cooling mass flow rate improves the cooling performance. When compared to the normal C3X turbine blade, the variation with the largest mass flow rate ṁ + 20% reduces the blade surface temperature by 2.21% which is equivalent to 14.48 K, increases the vane surface heat transfer coefficient by 23.55% which is equivalent to 87.22 W/m2K and increases the total heat transfer rate by 8.13% which is equivalent to 337.3 W.