I have 2 questions. 1. The problem statement, all variables and given/known data 1)An aircraft is flying at sea level at a speed of 280km/h. Calculate the static pressure and total pressure at the stagnation point. From a standard atmosphere table, for sea level altitude, pressure is 101325 Pa. Density of air is 1.2250 kg/m3. 2) consider an airplane flying with a velocity of 60m/s at a standard altitude of 3km. At a point on the wing, the airflow velocity is 70m/s. Calculate the pressure at this point. Assume an incompressible flow. From a standard atmosphere table, for altitude of 3km, pressure = 70121 Pa, density of air is 0.90926 kg/m3. 2. Relevant equations for question 2, what is exactly meant by a standard altitude of 3km? does it mean that the airplane is in fact flying at 3km above sea level? or does it mean that the airplane is flying at a certain altitude and speed such that it experiences the pressure that a stationary object at 3km would experience? 3. The attempt at a solution For question 1, i think that the static pressure at stagnation point should be equal to the atmosphere pressure which is 101325 Pa. The total pressure would be static pressure + dynamic pressure which is 101325 + (1/2)(1.2250)(280 000/3600)2 = 105030 Pa For question 2, Assuming that the airplane is indeed flying at 3km above sea level, let p0 be the pressure at a point far ahead in the free stream where the velocity of air is 0m/s. Let p1 be the pressure at the point of the wing where velocity of air is 70m/s. then using p0 + (1/2)*rho*v02 = p1 + (1/2)*rho*v12, 70121 + 0 = p1 + (1/2)*(0.90926)*(70)2 p1 = 67893 Pa Thanks for any help rendered..