XFOIL Version 6.96 Calculated polar for: SPICA 11.73% smoothed 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.500 -0.3240 0.13326 0.13151 -0.0287 1.0000 0.0159 -12.250 -0.3250 0.12914 0.12740 -0.0309 1.0000 0.0160 -12.000 -0.3227 0.12555 0.12382 -0.0323 1.0000 0.0160 -11.750 -0.3173 0.12182 0.12010 -0.0323 1.0000 0.0162 -11.500 -0.3113 0.11904 0.11733 -0.0327 1.0000 0.0163 -11.250 -0.3035 0.11692 0.11522 -0.0326 1.0000 0.0165 -11.000 -0.2983 0.11427 0.11259 -0.0332 1.0000 0.0166 -10.750 -0.2931 0.11176 0.11009 -0.0336 1.0000 0.0168 -10.500 -0.2885 0.10916 0.10750 -0.0341 1.0000 0.0171 -10.250 -0.1817 0.09217 0.09062 -0.0500 0.9731 0.0193 -10.000 -0.1669 0.08581 0.08423 -0.0565 0.9605 0.0194 -9.750 -0.1419 0.07873 0.07709 -0.0643 0.9456 0.0197 -9.500 -0.1029 0.07299 0.07123 -0.0736 0.9148 0.0199 -9.250 -0.0895 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0.0120 1.0000 15.250 1.4971 0.05060 0.04496 -0.0247 0.0117 1.0000 15.500 1.4956 0.05356 0.04799 -0.0246 0.0114 1.0000 15.750 1.4905 0.05699 0.05151 -0.0246 0.0112 1.0000 16.000 1.4859 0.06044 0.05505 -0.0248 0.0110 1.0000 16.250 1.4783 0.06435 0.05906 -0.0251 0.0109 1.0000 16.500 1.4670 0.06879 0.06359 -0.0256 0.0106 1.0000 16.750 1.4523 0.07382 0.06874 -0.0263 0.0104 1.0000 17.000 1.4465 0.07774 0.07275 -0.0269 0.0103 1.0000 17.250 1.4414 0.08160 0.07670 -0.0276 0.0102 1.0000 17.500 1.4332 0.08596 0.08116 -0.0285 0.0102 1.0000 17.750 1.4265 0.09020 0.08549 -0.0294 0.0100 1.0000 18.000 1.4189 0.09461 0.08999 -0.0305 0.0099 1.0000 18.250 1.4086 0.09945 0.09493 -0.0318 0.0099 1.0000 18.500 1.3985 0.10437 0.09994 -0.0332 0.0097 1.0000 18.750 1.3923 0.10875 0.10440 -0.0345 0.0095 1.0000