Consider a low-speed, steady flow around the thin airfoil shown in Fig. P3.3. We know the velocity and altitude at which the vehicle is flying. Thus, we know p. (i.e., p.) and U.. We have obtained experimental values of the local static pressure at points 2 through 6. At which of these points can we use Bernoulli's equation to determine the local velocity? If we cannot, why not? bint 1 3034 Edge of boundary layer 5 Point 2: at the stagnation point of the airfoil at a point in the inviscid region just outside the laminar boundary layer at a point in the laminar boundary layer Point 3: Point 4: Point 5: at a point in the turbulent boundary layer Point 6: at a point in the inviscid region just outside the turbulent boundary layer Assume that the airfoil of problem 3.3 is moving at 300 km/h at an altitude of 3 km. The experimentally determined pressure coefficients are: Point 2 3 C, 1.00 -3.00 -3.00 5 6 +0.16 +0.16 (a) What is the Mach number and the Reynolds number for this configuration? Assume that the characteristic dimension for the airfoil is 1.5 m. (b) Calculate the local pressure in N/m² and in lbf/in² at all five points. What is the percentage change in the pressure relative to the free-stream value? That is, what is (Plocal p.)/p.? Was it reasonable to assume that the pressure changes are sufficiently small that the density is approximately constant? (c) Why are the pressures at points 3 and 4 equal and at points 5 and 6 equal? (d) At those points where Bernoulli's equation can be used validly, calculate the local velocity.