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Fluid dynamics of an aircraft

  1. Nov 12, 2008 #1
    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..
  2. jcsd
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