XFOIL Version 6.94 Calculated polar for: AG45c -03f 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0888 0.00665 0.00207 -0.0298 0.7352 1.0000 0.500 0.1434 0.00699 0.00199 -0.0292 0.7069 1.0000 1.000 0.1989 0.00719 0.00195 -0.0288 0.6781 1.0000 1.500 0.2546 0.00729 0.00186 -0.0284 0.6481 1.0000 2.000 0.3102 0.00741 0.00181 -0.0280 0.6151 1.0000 2.500 0.3655 0.00761 0.00180 -0.0276 0.5803 1.0000 3.000 0.4209 0.00785 0.00188 -0.0272 0.5431 1.0000 3.500 0.4762 0.00814 0.00203 -0.0269 0.5045 1.0000 4.000 0.5313 0.00848 0.00223 -0.0266 0.4653 1.0000 4.500 0.5863 0.00886 0.00251 -0.0263 0.4238 1.0000 5.000 0.6412 0.00929 0.00283 -0.0261 0.3789 1.0000 5.500 0.6955 0.00984 0.00327 -0.0258 0.3312 1.0000 6.000 0.7496 0.01047 0.00379 -0.0256 0.2797 1.0000 6.500 0.8030 0.01124 0.00446 -0.0254 0.2273 1.0000 7.000 0.8556 0.01217 0.00527 -0.0252 0.1787 1.0000 7.500 0.9078 0.01321 0.00627 -0.0249 0.1363 1.0000 8.000 0.9588 0.01444 0.00742 -0.0246 0.0986 1.0000 8.500 1.0088 0.01584 0.00881 -0.0241 0.0653 1.0000 9.000 1.0565 0.01774 0.01077 -0.0234 0.0418 1.0000 9.500 1.1007 0.02022 0.01347 -0.0223 0.0294 1.0000 10.000 1.1412 0.02323 0.01678 -0.0209 0.0225 1.0000 10.500 1.1759 0.02711 0.02094 -0.0192 0.0185 1.0000 11.000 1.2017 0.03277 0.02723 -0.0169 0.0165 1.0000 11.500 1.2278 0.03679 0.03178 -0.0153 0.0143 1.0000 12.000 1.2337 0.04282 0.03837 -0.0135 0.0134 1.0000 12.500 1.2161 0.05021 0.04624 -0.0138 0.0130 1.0000 13.000 1.1880 0.06164 0.05814 -0.0210 0.0130 1.0000 13.500 1.1499 0.07673 0.07363 -0.0314 0.0133 1.0000 14.000 1.0965 0.09696 0.09418 -0.0449 0.0140 1.0000 14.500 1.0206 0.12576 0.12320 -0.0618 0.0158 1.0000 15.000 0.7681 0.14729 0.14475 -0.0522 0.0227 1.0000 15.500 0.7405 0.15777 0.15520 -0.0558 0.0241 1.0000