XFOIL Version 6.96 Calculated polar for: WORTMANN FX 2 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.5210 0.09132 0.08647 -0.1598 0.8226 0.0093 -19.500 -0.5337 0.08567 0.08071 -0.1629 0.8224 0.0094 -19.250 -0.5449 0.08048 0.07541 -0.1657 0.8221 0.0094 -19.000 -0.5538 0.07576 0.07057 -0.1682 0.8218 0.0094 -18.750 -0.5605 0.07145 0.06616 -0.1706 0.8215 0.0096 -18.500 -0.5660 0.06749 0.06211 -0.1726 0.8212 0.0097 -18.250 -0.5663 0.06430 0.05883 -0.1743 0.8209 0.0098 -18.000 -0.5690 0.06095 0.05538 -0.1758 0.8205 0.0099 -17.750 -0.5700 0.05793 0.05228 -0.1771 0.8202 0.0100 -17.500 -0.5690 0.05523 0.04949 -0.1781 0.8198 0.0103 -17.250 -0.5670 0.05276 0.04692 -0.1790 0.8195 0.0104 -17.000 -0.5628 0.05057 0.04464 -0.1797 0.8191 0.0107 -16.750 -0.5605 0.04829 0.04227 -0.1803 0.8187 0.0108 -16.500 -0.5552 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0.04340 -0.0253 0.0136 0.7226 14.500 1.1663 0.05251 0.04499 -0.0245 0.0132 0.7231 14.750 1.1771 0.05398 0.04653 -0.0239 0.0127 0.7237 15.000 1.1867 0.05558 0.04819 -0.0232 0.0122 0.7241 15.250 1.1965 0.05718 0.04987 -0.0227 0.0119 0.7245 15.500 1.2044 0.05900 0.05176 -0.0221 0.0114 0.7249 15.750 1.2122 0.06086 0.05368 -0.0215 0.0110 0.7252 16.000 1.2178 0.06298 0.05586 -0.0210 0.0106 0.7256 16.250 1.2248 0.06499 0.05794 -0.0206 0.0102 0.7259 16.500 1.2302 0.06723 0.06027 -0.0202 0.0100 0.7262 16.750 1.2372 0.06932 0.06246 -0.0199 0.0097 0.7266 17.250 1.2474 0.07400 0.06732 -0.0196 0.0092 0.7274 17.500 1.2519 0.07646 0.06987 -0.0195 0.0089 0.7278 17.750 1.2559 0.07905 0.07255 -0.0196 0.0087 0.7282 18.000 1.2588 0.08180 0.07539 -0.0197 0.0086 0.7287 18.250 1.2600 0.08480 0.07849 -0.0200 0.0084 0.7291 18.500 1.2603 0.08800 0.08178 -0.0203 0.0083 0.7295 18.750 1.2576 0.09163 0.08550 -0.0209 0.0080 0.7299