XFOIL Version 6.96 Calculated polar for: FX 79-L-100 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.8806 0.09620 0.09383 0.0085 1.0000 0.0146 -13.250 -0.8907 0.08797 0.08548 0.0029 1.0000 0.0144 -13.000 -0.9333 0.07434 0.07172 -0.0067 1.0000 0.0140 -12.750 -0.9665 0.06409 0.06127 -0.0143 1.0000 0.0139 -12.500 -0.9848 0.05760 0.05458 -0.0190 1.0000 0.0140 -12.250 -0.9996 0.05237 0.04914 -0.0224 1.0000 0.0141 -12.000 -1.0229 0.04657 0.04310 -0.0246 1.0000 0.0139 -11.750 -1.0342 0.04308 0.03937 -0.0249 1.0000 0.0141 -11.500 -1.0420 0.04033 0.03639 -0.0237 1.0000 0.0142 -11.250 -1.0544 0.03658 0.03233 -0.0211 1.0000 0.0141 -11.000 -1.0496 0.03444 0.02994 -0.0195 1.0000 0.0143 -10.750 -1.0403 0.03267 0.02793 -0.0181 1.0000 0.0144 -10.500 -1.0433 0.02798 0.02282 -0.0157 1.0000 0.0150 -10.250 -1.0281 0.02653 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0.0187 1.0000 9.250 0.9522 0.02258 0.01689 0.0107 0.0181 1.0000 9.500 0.9718 0.02367 0.01812 0.0118 0.0173 1.0000 9.750 0.9908 0.02471 0.01929 0.0128 0.0167 1.0000 10.000 1.0084 0.02597 0.02068 0.0139 0.0161 1.0000 10.250 1.0273 0.02666 0.02141 0.0147 0.0155 1.0000 10.500 1.0423 0.02817 0.02304 0.0159 0.0151 1.0000 10.750 1.0506 0.03083 0.02593 0.0175 0.0148 1.0000 11.250 1.0502 0.03727 0.03305 0.0213 0.0142 1.0000 11.500 1.0430 0.04014 0.03621 0.0237 0.0141 1.0000 11.750 1.0343 0.04304 0.03933 0.0249 0.0141 1.0000 12.000 1.0199 0.04704 0.04357 0.0245 0.0141 1.0000 12.250 1.0012 0.05218 0.04893 0.0224 0.0141 1.0000 12.500 0.9793 0.05850 0.05548 0.0186 0.0142 1.0000 12.750 0.9573 0.06570 0.06288 0.0134 0.0142 1.0000 13.000 0.9228 0.07614 0.07348 0.0059 0.0144 1.0000