However for a flap chord ratio of 0.1, the predicted values of C,iiiai based on Figures 3.32 to 3.34 are higher than those shown in Figure 3.26a. If the procedure is repeated for other flap angles, close agreement is also obtained with the figure. This result compares closely with Figure 3.26a. Thus, for the flapped airfoil, C(nm is predicted to be 1.65+0.96, or 2.61. Hence,Īccording to Figure 3.27, Cimn for a plain 23012 airfoil equals 1.65 at R = 3.5 x 106. Using Figure 3.34, the ratio of ACJmar to AC, is obtained as 0.66. In Equation 3.51, Cta of 0.015 is obtained from Reference 3.1. Hence from Equation 3.49, AC/ is equal to From Figure 3.32, t - 0.66 for cflc = 0.3 and from Figure 3.33, v = 0.35 for a split flap deflected 60°. Figure 3.34 Gmax increment ratio as a function of flap chord ratio.Īs an example, in using Figures 3.32, 3.33, and 3.34, consider the prediction of Ctwix for a 23012 airfoil equipped with a 30% chord split flap deflected 60° and operating at a Reynold's number of 3.5 x 10*.
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