XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.250 -0.7825 0.08704 0.08170 -0.0911 1.0000 0.0503 -17.000 -0.8116 0.08006 0.07451 -0.0945 1.0000 0.0502 -16.750 -0.8384 0.07405 0.06827 -0.0970 1.0000 0.0501 -16.500 -0.8628 0.06898 0.06297 -0.0985 1.0000 0.0501 -16.250 -0.8852 0.06471 0.05849 -0.0991 1.0000 0.0502 -16.000 -0.9098 0.06182 0.05544 -0.0976 1.0000 0.0502 -15.750 -0.9449 0.06154 0.05514 -0.0912 1.0000 0.0502 -15.500 -0.9511 0.05767 0.05100 -0.0929 0.9962 0.0505 -15.250 -0.9280 0.05529 0.04860 -0.0950 0.9938 0.0511 -15.000 -0.9084 0.05306 0.04631 -0.0972 0.9913 0.0518 -14.750 -0.8902 0.05101 0.04418 -0.0991 0.9873 0.0525 -14.500 -0.8749 0.04865 0.04166 -0.1012 0.9836 0.0535 -14.250 -0.8612 0.04608 0.03886 -0.1034 0.9805 0.0545 -14.000 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0.2120 0.8482 10.000 1.1564 0.03631 0.02845 -0.0447 0.2030 0.8506 10.250 1.1712 0.03735 0.02956 -0.0441 0.1943 0.8533 10.500 1.1842 0.03857 0.03065 -0.0433 0.1869 0.8560 10.750 1.2006 0.03967 0.03182 -0.0431 0.1798 0.8589 11.000 1.2155 0.04090 0.03295 -0.0427 0.1735 0.8614 11.250 1.2288 0.04189 0.03400 -0.0416 0.1680 0.8639 11.500 1.2417 0.04296 0.03510 -0.0407 0.1628 0.8667 11.750 1.2570 0.04394 0.03595 -0.0398 0.1579 0.8700 12.000 1.2712 0.04514 0.03729 -0.0393 0.1535 0.8736 12.250 1.2865 0.04638 0.03854 -0.0391 0.1492 0.8771 12.500 1.3038 0.04719 0.03922 -0.0384 0.1450 0.8800 12.750 1.3144 0.04849 0.04068 -0.0375 0.1416 0.8832 13.000 1.3275 0.04970 0.04195 -0.0368 0.1380 0.8867 13.250 1.3426 0.05085 0.04308 -0.0364 0.1347 0.8903 13.500 1.3613 0.05190 0.04410 -0.0363 0.1314 0.8938 13.750 1.3703 0.05333 0.04568 -0.0353 0.1284 0.8973 14.000 1.3815 0.05463 0.04704 -0.0346 0.1255 0.9013 14.250 1.3967 0.05572 0.04809 -0.0341 0.1228 0.9058 14.500 1.4134 0.05686 0.04924 -0.0337 0.1201 0.9104 14.750 1.4196 0.05848 0.05104 -0.0327 0.1177 0.9153 15.000 1.4293 0.06002 0.05266 -0.0320 0.1153 0.9207 15.250 1.4424 0.06129 0.05391 -0.0316 0.1130 0.9259 15.500 1.4619 0.06180 0.05432 -0.0308 0.1105 0.9318 15.750 1.4637 0.06411 0.05687 -0.0302 0.1087 0.9381 16.000 1.4674 0.06597 0.05889 -0.0292 0.1068 0.9456 16.250 1.4742 0.06760 0.06060 -0.0284 0.1049 0.9553 16.500 1.4830 0.06871 0.06171 -0.0275 0.1031 0.9748 16.750 1.5073 0.06915 0.06203 -0.0277 0.1010 1.0000 17.000 1.5060 0.07243 0.06557 -0.0282 0.0995 1.0000 17.250 1.5077 0.07551 0.06884 -0.0287 0.0978 1.0000 17.500 1.5139 0.07812 0.07156 -0.0294 0.0962 1.0000 17.750 1.5244 0.08023 0.07370 -0.0300 0.0946 1.0000 18.000 1.5446 0.08125 0.07463 -0.0305 0.0929 1.0000 18.250 1.5516 0.08376 0.07724 -0.0310 0.0915 1.0000 18.500 1.5429 0.08799 0.08173 -0.0318 0.0903 1.0000 18.750 1.5373 0.09198 0.08592 -0.0328 0.0891 1.0000 19.000 1.5337 0.09578 0.08987 -0.0339 0.0878 1.0000 19.250 1.5352 0.09893 0.09312 -0.0349 0.0866 1.0000