XFOIL Version 6.94 Calculated polar for: E224 (10.17%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2450 0.00618 0.00186 -0.0566 0.7948 0.8215 0.500 0.2948 0.00620 0.00189 -0.0546 0.7646 0.8701 1.000 0.3430 0.00626 0.00192 -0.0522 0.7327 0.9153 1.500 0.3978 0.00635 0.00195 -0.0513 0.6981 0.9556 2.000 0.4708 0.00653 0.00199 -0.0547 0.6581 0.9834 2.500 0.5462 0.00676 0.00207 -0.0590 0.6129 1.0000 3.000 0.5889 0.00706 0.00220 -0.0565 0.5694 1.0000 3.500 0.6355 0.00745 0.00242 -0.0546 0.5213 1.0000 4.000 0.6840 0.00789 0.00270 -0.0532 0.4699 1.0000 4.500 0.7321 0.00847 0.00306 -0.0518 0.4083 1.0000 5.000 0.7810 0.00908 0.00349 -0.0506 0.3511 1.0000 5.500 0.8280 0.00990 0.00403 -0.0492 0.2802 1.0000 6.000 0.8736 0.01092 0.00472 -0.0477 0.2023 1.0000 6.500 0.9197 0.01193 0.00549 -0.0463 0.1431 1.0000 7.000 0.9646 0.01305 0.00640 -0.0447 0.0948 1.0000 7.500 1.0069 0.01442 0.00759 -0.0428 0.0517 1.0000 8.000 1.0469 0.01602 0.00915 -0.0402 0.0305 1.0000 9.000 1.1225 0.01918 0.01261 -0.0348 0.0203 1.0000 10.500 1.1818 0.02733 0.02144 -0.0201 0.0120 1.0000 11.000 1.2035 0.02921 0.02357 -0.0160 0.0106 1.0000 11.500 1.2188 0.03158 0.02612 -0.0122 0.0093 1.0000 12.000 1.2240 0.03508 0.02983 -0.0086 0.0083 1.0000 12.500 1.2100 0.04154 0.03670 -0.0053 0.0079 1.0000 13.000 1.1857 0.04985 0.04549 -0.0041 0.0076 1.0000 13.500 1.1667 0.05821 0.05424 -0.0056 0.0076 1.0000 14.000 1.1482 0.06757 0.06395 -0.0096 0.0074 1.0000 14.500 1.1156 0.08078 0.07753 -0.0165 0.0074 1.0000 15.000 1.0753 0.09747 0.09455 -0.0265 0.0075 1.0000 15.500 1.0312 0.11726 0.11465 -0.0387 0.0076 1.0000