XFOIL Version 6.96 Calculated polar for: WORTMANN FX 79-K-144/17 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.750 -0.6074 0.06342 0.05919 -0.0713 0.8249 0.0141 -11.500 -0.6330 0.05620 0.05175 -0.0763 0.8241 0.0140 -11.250 -0.6589 0.05041 0.04566 -0.0780 0.8234 0.0140 -11.000 -0.6817 0.04537 0.04023 -0.0775 0.8227 0.0141 -10.750 -0.6845 0.04300 0.03769 -0.0765 0.8220 0.0142 -10.500 -0.6816 0.04151 0.03606 -0.0751 0.8214 0.0145 -10.250 -0.6724 0.04015 0.03456 -0.0743 0.8208 0.0149 -10.000 -0.6627 0.03862 0.03283 -0.0734 0.8200 0.0154 -9.750 -0.6536 0.03663 0.03053 -0.0723 0.8190 0.0162 -9.500 -0.6444 0.03394 0.02738 -0.0711 0.8177 0.0173 -9.250 -0.6316 0.03142 0.02436 -0.0699 0.8163 0.0183 -9.000 -0.6128 0.03020 0.02307 -0.0696 0.8148 0.0190 -8.750 -0.5926 0.02916 0.02188 -0.0692 0.8134 0.0200 -8.500 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0.04034 0.03246 -0.0206 0.0175 1.0000 11.750 1.1166 0.04212 0.03433 -0.0199 0.0164 1.0000 12.000 1.1245 0.04396 0.03626 -0.0193 0.0157 1.0000 12.250 1.1315 0.04593 0.03830 -0.0186 0.0151 1.0000 12.500 1.1377 0.04801 0.04046 -0.0180 0.0146 1.0000 12.750 1.1457 0.04999 0.04257 -0.0174 0.0141 1.0000 13.000 1.1537 0.05197 0.04467 -0.0169 0.0136 1.0000 13.250 1.1619 0.05398 0.04679 -0.0165 0.0130 1.0000 13.500 1.1698 0.05605 0.04896 -0.0162 0.0125 1.0000 13.750 1.1772 0.05820 0.05122 -0.0161 0.0120 1.0000 14.000 1.1836 0.06049 0.05360 -0.0160 0.0117 1.0000 14.250 1.1885 0.06294 0.05613 -0.0158 0.0115 1.0000 14.500 1.1933 0.06545 0.05876 -0.0155 0.0113 1.0000 14.750 1.1978 0.06803 0.06150 -0.0152 0.0111 1.0000 15.000 1.2014 0.07076 0.06439 -0.0149 0.0109 1.0000 15.250 1.2038 0.07369 0.06750 -0.0148 0.0107 1.0000 15.500 1.2048 0.07687 0.07086 -0.0148 0.0106 1.0000 15.750 1.2043 0.08031 0.07449 -0.0149 0.0105 1.0000 16.000 1.2020 0.08408 0.07846 -0.0154 0.0104 1.0000 16.250 1.1978 0.08822 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