In film cooling of a gas turbine blade, the cooling air is injected through a series of cooling holes into the main stream to form an insulating blanket over the blade surface. However, it has to pay the aerodynamic penalties due to the coolant injection, which are subject to the impact of many geometric and flow conditions. Because film cooling with backward injection is reported to be able to provide a better cooling performance, this paper studies the losses occurred on the turbine blade with backward injection and compares it to the case with forward injection. The pressure loss coefficient is used in this study to report the aerodynamic loss. Different blowing ratios are considered. The numerical simulation indicates that the loss is higher with backward injection than that with forward injection for all the blowing ratios under study. To compare with the cylindrical holes, fan-shaped cooling holes are also considered. It is observed that the loss is larger for the fan-shaped holes in the forward injection whereas there is no substantial difference in the case of backward injection. The aerodynamic loss due to the location of coolant injection is studied by using injection from the leading edge, pressure side, suction side and trailing edge, respectively. The effects of cooling air injection angles and main stream incidence angles on the aerodynamic loss are also reported.