XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3782 0.05215 0.04112 -0.0075 0.9999 0.1928 -2.750 -0.3517 0.04923 0.03777 -0.0073 0.9999 0.1841 -2.500 -0.3226 0.04691 0.03483 -0.0069 0.9999 0.1772 -2.250 -0.2918 0.04446 0.03206 -0.0067 0.9999 0.1732 -2.000 -0.2576 0.04228 0.02948 -0.0069 0.9999 0.1698 -1.750 -0.2209 0.04032 0.02716 -0.0072 0.9999 0.1682 -1.500 -0.1829 0.03855 0.02517 -0.0077 0.9999 0.1690 -1.250 -0.1442 0.03695 0.02351 -0.0082 0.9999 0.1728 -1.000 -0.1065 0.03549 0.02216 -0.0084 0.9999 0.1803 -0.750 -0.0706 0.03402 0.02109 -0.0086 0.9999 0.1965 -0.500 -0.0340 0.03227 0.02023 -0.0091 0.9999 0.2373 -0.250 0.1264 0.02797 0.01801 -0.0307 0.8356 1.0001 0.000 0.1745 0.02906 0.01716 -0.0311 0.5548 1.0001 0.250 0.1998 0.03004 0.01743 -0.0297 0.5306 1.0001 0.500 0.2273 0.03075 0.01776 -0.0289 0.5166 1.0001 0.750 0.2554 0.03147 0.01811 -0.0283 0.5086 1.0001 1.000 0.2846 0.03202 0.01850 -0.0280 0.5019 1.0001 1.250 0.3062 0.03309 0.01903 -0.0263 0.4680 1.0001 1.500 0.3236 0.03413 0.01947 -0.0240 0.4184 1.0001 1.750 0.3429 0.03494 0.01978 -0.0222 0.3822 1.0001 2.000 0.3647 0.03543 0.01999 -0.0209 0.3575 1.0001 2.250 0.3867 0.03583 0.02014 -0.0197 0.3377 1.0001 2.500 0.4095 0.03606 0.02023 -0.0185 0.3207 1.0001 2.750 0.4322 0.03626 0.02027 -0.0173 0.3038 1.0001 3.000 0.4548 0.03650 0.02024 -0.0161 0.2856 1.0001 3.250 0.4777 0.03681 0.02040 -0.0149 0.2693 1.0001 3.500 0.5018 0.03721 0.02073 -0.0138 0.2559 1.0001 3.750 0.5263 0.03767 0.02122 -0.0128 0.2384 1.0001 4.000 0.5521 0.03838 0.02201 -0.0118 0.2235 1.0001 4.250 0.5785 0.03938 0.02310 -0.0109 0.2067 1.0001 4.500 0.6064 0.04082 0.02467 -0.0101 0.1882 1.0001 4.750 0.6352 0.04272 0.02680 -0.0094 0.1672 1.0001 5.000 0.6652 0.04517 0.02947 -0.0091 0.1539 1.0001