XFOIL Version 6.96 Calculated polar for: WORTMANN FX 71-120 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.6706 0.06641 0.05793 -0.0140 1.0000 0.2317 -7.500 -0.6702 0.06222 0.05369 -0.0135 1.0000 0.2285 -7.250 -0.6835 0.05716 0.04845 -0.0128 1.0000 0.2241 -7.000 -0.7089 0.05123 0.04198 -0.0111 1.0000 0.2185 -6.750 -0.7159 0.04752 0.03767 -0.0088 1.0000 0.2161 -6.500 -0.7054 0.04477 0.03469 -0.0071 1.0000 0.2158 -6.250 -0.6934 0.04220 0.03190 -0.0054 1.0000 0.2156 -6.000 -0.6804 0.03986 0.02929 -0.0037 1.0000 0.2158 -5.750 -0.6658 0.03769 0.02687 -0.0020 1.0000 0.2164 -5.500 -0.6477 0.03570 0.02481 -0.0007 1.0000 0.2181 -5.250 -0.6287 0.03401 0.02305 0.0007 1.0000 0.2203 -5.000 -0.6092 0.03253 0.02150 0.0020 1.0000 0.2238 -4.750 -0.5902 0.03114 0.01995 0.0035 1.0000 0.2287 -4.500 -0.5715 0.02980 0.01833 0.0051 1.0000 0.2340 -4.250 -0.5495 0.02851 0.01721 0.0062 1.0000 0.2403 -4.000 -0.5285 0.02736 0.01594 0.0076 1.0000 0.2486 -3.750 -0.5060 0.02616 0.01486 0.0087 1.0000 0.2580 -3.500 -0.4837 0.02507 0.01380 0.0100 1.0000 0.2723 -3.250 -0.4617 0.02402 0.01297 0.0113 1.0000 0.2930 -3.000 -0.4410 0.02294 0.01216 0.0128 1.0000 0.3276 -2.750 -0.4224 0.02170 0.01150 0.0147 1.0000 0.3875 -2.500 -0.4104 0.02048 0.01112 0.0182 1.0000 0.4873 -2.250 -0.4021 0.01978 0.01122 0.0232 1.0000 0.6018 -2.000 -0.3899 0.01971 0.01172 0.0286 1.0000 0.7054 -1.750 -0.3696 0.02015 0.01244 0.0332 1.0000 0.7945 -1.500 -0.3052 0.02163 0.01392 0.0316 1.0000 0.8764 -1.250 -0.1115 0.02355 0.01540 0.0075 1.0000 0.9498 -1.000 0.0290 0.02258 0.01419 -0.0135 1.0000 0.9965 -0.750 0.0368 0.02205 0.01368 -0.0128 1.0000 1.0000 -0.500 0.0282 0.02177 0.01344 -0.0091 1.0000 1.0000 -0.250 0.0153 0.02162 0.01331 -0.0047 1.0000 1.0000 0.000 0.0000 0.02157 0.01327 0.0000 1.0000 1.0000 0.250 -0.0153 0.02162 0.01331 0.0047 1.0000 1.0000 0.500 -0.0281 0.02177 0.01343 0.0091 1.0000 1.0000 0.750 -0.0368 0.02205 0.01367 0.0128 1.0000 1.0000 1.000 -0.0292 0.02257 0.01418 0.0135 0.9966 1.0000 1.250 0.1113 0.02354 0.01539 -0.0074 0.9498 1.0000 1.500 0.3051 0.02162 0.01392 -0.0316 0.8764 1.0000 1.750 0.3696 0.02015 0.01244 -0.0332 0.7947 1.0000 2.000 0.3899 0.01970 0.01172 -0.0286 0.7054 1.0000 2.250 0.4021 0.01978 0.01122 -0.0232 0.6020 1.0000 2.500 0.4103 0.02048 0.01112 -0.0182 0.4867 1.0000 2.750 0.4223 0.02169 0.01150 -0.0147 0.3874 1.0000 3.000 0.4410 0.02294 0.01217 -0.0128 0.3278 1.0000 3.250 0.4617 0.02402 0.01296 -0.0113 0.2932 1.0000 3.500 0.4837 0.02507 0.01380 -0.0100 0.2725 1.0000 3.750 0.5059 0.02616 0.01485 -0.0087 0.2582 1.0000 4.000 0.5284 0.02735 0.01594 -0.0076 0.2486 1.0000 4.250 0.5496 0.02851 0.01721 -0.0062 0.2404 1.0000 4.500 0.5715 0.02981 0.01834 -0.0051 0.2339 1.0000 4.750 0.5902 0.03114 0.01995 -0.0035 0.2286 1.0000 5.000 0.6091 0.03253 0.02149 -0.0020 0.2237 1.0000 5.250 0.6287 0.03401 0.02304 -0.0007 0.2203 1.0000 5.500 0.6477 0.03570 0.02480 0.0007 0.2180 1.0000 5.750 0.6658 0.03769 0.02687 0.0020 0.2164 1.0000 6.000 0.6804 0.03986 0.02929 0.0037 0.2158 1.0000 6.250 0.6934 0.04221 0.03190 0.0054 0.2156 1.0000 6.500 0.7055 0.04476 0.03468 0.0071 0.2158 1.0000 6.750 0.7160 0.04752 0.03767 0.0087 0.2161 1.0000 7.000 0.7092 0.05121 0.04195 0.0111 0.2184 1.0000 7.250 0.6836 0.05716 0.04845 0.0128 0.2241 1.0000 7.500 0.6705 0.06223 0.05369 0.0135 0.2286 1.0000 7.750 0.6708 0.06641 0.05792 0.0139 0.2317 1.0000