Discuss the difference in results for both airfoils.
The NACA 653-018 airfoil (Table P6.1) is an example of a low drag or laminar flow airfoil. The subscript 3 indicates that the drag coefficient is a minimum over a range of lift coefficients of 0.3 on either side of the design lift coefficient, which for this airfoil section is zero. The performance characteristics of this airfoil differ from a conventional airfoil, like the NACA 0012 airfoil, in that near the design lift coefficient, the low-drag airfoil has a laminar boundary layer over a considerable portion of its area. The late transition is accomplished by designing the airfoil so that the pressure gradient is favorable over a larger area. A favorable pressure gradient tends to delay transition. The main geometrical difference that accounts for the change in the pressure distribution is a movement rearward of the point of maximum thickness. (a) Repeat Problem 6.4 for this airfoil and compare your calculated results with the experimental data (Fig. P6.2) obtained for a chord Reynolds number of 6 x 106. (b) Discuss the difference in results for both airfoils. Problem 6.4 The NACA 0012 airfoil is a conventional airfoil which has a favorable pressure distribution on the upper surface up to about a quarter chord point at a = 0°; with increasing incidence angle, say a = 8°, the gradient becomes unfavorable over practically the entire surface. (a) Compute the pressure distribution on this airfoil with the panel program of Section 6.5 for angles of attack of a = 0°, 4°, 10° and plot Cp vs x/c and V/V∞ vs x/c for each α. (b) Compute the lift coefficients for angles of attack of a from 0° to 20° at 4° degree intervals and compare them with the experimental data in Fig. P6.1 obtained for a chord Reynolds number Rc(≡ V∞C/V) of 6 x 106. Discuss the numerical results with experimental data.
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