XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421MOD 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 ------ -------- --------- --------- -------- -------- -------- -19.500 -0.8560 0.11804 0.11233 -0.0399 1.0000 0.0322 -19.250 -0.8796 0.10961 0.10366 -0.0449 1.0000 0.0324 -19.000 -0.8964 0.10272 0.09655 -0.0488 1.0000 0.0326 -18.750 -0.9094 0.09665 0.09029 -0.0521 1.0000 0.0328 -18.500 -0.9135 0.09229 0.08582 -0.0543 1.0000 0.0330 -18.250 -0.9131 0.08875 0.08224 -0.0559 1.0000 0.0332 -18.000 -0.9135 0.08511 0.07855 -0.0576 1.0000 0.0334 -17.750 -0.9135 0.08156 0.07493 -0.0591 1.0000 0.0337 -17.500 -0.9133 0.07809 0.07139 -0.0606 1.0000 0.0340 -17.250 -0.9128 0.07471 0.06794 -0.0621 1.0000 0.0343 -17.000 -0.9118 0.07146 0.06460 -0.0633 1.0000 0.0346 -16.750 -0.9100 0.06838 0.06142 -0.0645 1.0000 0.0350 -16.500 -0.9076 0.06544 0.05839 -0.0655 1.0000 0.0355 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0.04482 -0.0513 0.1415 0.8163 12.750 1.4107 0.05452 0.04682 -0.0509 0.1387 0.8190 13.000 1.4167 0.05668 0.04906 -0.0507 0.1360 0.8218 13.250 1.4216 0.05903 0.05146 -0.0506 0.1334 0.8246 13.500 1.4252 0.06159 0.05404 -0.0506 0.1312 0.8275 13.750 1.4313 0.06394 0.05645 -0.0507 0.1291 0.8304 14.000 1.4377 0.06622 0.05883 -0.0507 0.1269 0.8330 14.250 1.4422 0.06862 0.06132 -0.0506 0.1247 0.8355 14.500 1.4454 0.07123 0.06400 -0.0507 0.1227 0.8384 14.750 1.4483 0.07392 0.06672 -0.0508 0.1208 0.8415 15.000 1.4515 0.07663 0.06947 -0.0511 0.1190 0.8447 15.250 1.4570 0.07921 0.07216 -0.0515 0.1170 0.8479 15.500 1.4607 0.08199 0.07503 -0.0520 0.1151 0.8506 15.750 1.4634 0.08474 0.07787 -0.0522 0.1133 0.8530 16.000 1.4654 0.08763 0.08082 -0.0527 0.1117 0.8557 16.250 1.4681 0.09042 0.08363 -0.0531 0.1102 0.8588 16.500 1.4718 0.09315 0.08641 -0.0536 0.1088 0.8622 16.750 1.4748 0.09614 0.08953 -0.0544 0.1072 0.8658 17.000 1.4772 0.09904 0.09253 -0.0550 0.1057 0.8687 17.250 1.4785 0.10209 0.09567 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