At a given point in the high-speed flow over an airplane wing, the local Mach number, pressure, and temperature are 0.7, \(0.9 \mathrm{~atm}\), and \(250 \mathrm{~K}\), respectively. Calculate the values of \(p_{o}, T_{o}, p^{*}, T^{*}\), and \(a^{*}\) at this point.