XFOIL Version 6.94 Calculated polar for: NASA/LANGLEY MS(1)-0313 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3785 0.00859 0.00393 -0.0801 0.7852 0.7471 0.500 0.4365 0.00850 0.00377 -0.0802 0.7637 0.7540 1.000 0.4938 0.00849 0.00373 -0.0801 0.7423 0.7602 1.500 0.5513 0.00852 0.00371 -0.0802 0.7122 0.7671 2.000 0.6082 0.00869 0.00369 -0.0802 0.6577 0.7738 2.500 0.6551 0.00965 0.00385 -0.0785 0.4863 0.7799 3.000 0.6993 0.01114 0.00458 -0.0769 0.3186 0.7866 3.500 0.7482 0.01223 0.00516 -0.0761 0.2206 0.7937 4.000 0.8001 0.01307 0.00571 -0.0758 0.1712 0.7997 4.500 0.8507 0.01369 0.00624 -0.0750 0.1461 0.8070 5.000 0.9005 0.01446 0.00695 -0.0741 0.1278 0.8140 5.500 0.9523 0.01507 0.00761 -0.0736 0.1193 0.8212 6.000 1.0020 0.01577 0.00833 -0.0727 0.1115 0.8288 6.500 1.0474 0.01677 0.00938 -0.0710 0.1053 0.8365 7.000 1.0904 0.01825 0.01099 -0.0690 0.1010 0.8448 7.500 1.1369 0.01943 0.01237 -0.0677 0.0993 0.8533 8.000 1.1835 0.02001 0.01309 -0.0663 0.0979 0.8631 8.500 1.2305 0.02058 0.01377 -0.0651 0.0951 0.8731 9.000 1.2732 0.02135 0.01470 -0.0631 0.0918 0.8845 9.500 1.3139 0.02228 0.01576 -0.0609 0.0882 0.8981 10.000 1.3464 0.02314 0.01677 -0.0571 0.0847 0.9170 11.000 1.4074 0.02774 0.02185 -0.0505 0.0746 1.0000 11.500 1.4438 0.02812 0.02230 -0.0485 0.0734 1.0000 12.000 1.4782 0.02880 0.02309 -0.0465 0.0712 1.0000 12.500 1.5124 0.02964 0.02402 -0.0447 0.0670 1.0000 13.000 1.5467 0.03058 0.02500 -0.0431 0.0611 1.0000 13.500 1.5716 0.03258 0.02700 -0.0408 0.0547 1.0000 14.000 1.5856 0.03584 0.03047 -0.0378 0.0491 1.0000 14.500 1.6134 0.03782 0.03258 -0.0363 0.0460 1.0000 15.000 1.6324 0.04065 0.03542 -0.0347 0.0420 1.0000 15.500 1.6318 0.04558 0.04039 -0.0326 0.0375 1.0000 16.000 1.6381 0.05007 0.04516 -0.0312 0.0362 1.0000 16.500 1.6400 0.05535 0.05069 -0.0304 0.0346 1.0000 17.000 1.6343 0.06203 0.05760 -0.0306 0.0327 1.0000 17.500 1.6193 0.07053 0.06631 -0.0323 0.0309 1.0000 18.000 1.5917 0.08169 0.07772 -0.0361 0.0294 1.0000 18.500 1.5510 0.09597 0.09230 -0.0423 0.0282 1.0000 19.000 1.4972 0.11318 0.10985 -0.0508 0.0274 1.0000 19.500 1.4368 0.13239 0.12939 -0.0612 0.0267 1.0000 20.000 1.3758 0.15256 0.14986 -0.0731 0.0259 1.0000 20.500 1.3005 0.17781 0.17545 -0.0894 0.0248 1.0000 22.000 1.1201 0.26964 0.26761 -0.1484 0.0145 1.0000 22.500 1.1288 0.27836 0.27631 -0.1549 0.0121 1.0000 23.000 1.1431 0.28586 0.28382 -0.1602 0.0096 1.0000 24.000 1.1506 0.32320 0.32105 -0.1775 0.0124 1.0000 24.500 1.1658 0.33104 0.32891 -0.1829 0.0124 1.0000 25.000 1.1813 0.33825 0.33613 -0.1884 0.0123 1.0000 25.500 1.1967 0.34512 0.34300 -0.1939 0.0123 1.0000 26.000 1.2121 0.35163 0.34952 -0.1994 0.0123 1.0000 26.500 1.2274 0.35793 0.35583 -0.2049 0.0123 1.0000 27.000 1.2423 0.36403 0.36194 -0.2105 0.0123 1.0000 27.500 1.2571 0.36995 0.36787 -0.2161 0.0122 1.0000 28.000 1.2715 0.37565 0.37358 -0.2217 0.0122 1.0000 28.500 1.2857 0.38117 0.37914 -0.2274 0.0121 1.0000 29.000 1.2996 0.38651 0.38449 -0.2331 0.0121 1.0000 29.500 1.3133 0.39154 0.38953 -0.2389 0.0120 1.0000 30.000 1.3270 0.39576 0.39376 -0.2450 0.0117 1.0000 30.500 1.3424 0.39998 0.39800 -0.2516 0.0105 1.0000 31.000 1.3556 0.40534 0.40337 -0.2574 0.0092 1.0000 31.500 1.3676 0.41032 0.40838 -0.2630 0.0083 1.0000 32.000 1.3791 0.41503 0.41311 -0.2686 0.0075 1.0000 32.500 1.3899 0.41937 0.41747 -0.2741 0.0069 1.0000 33.000 1.3999 0.42316 0.42127 -0.2794 0.0063 1.0000 33.500 1.4094 0.42681 0.42497 -0.2846 0.0059 1.0000 34.000 1.4180 0.43011 0.42831 -0.2893 0.0055 1.0000 35.500 1.4516 0.45001 0.44820 -0.3087 0.0030 1.0000 36.000 1.4596 0.45400 0.45222 -0.3143 0.0026 1.0000 36.500 1.4664 0.45736 0.45559 -0.3198 0.0023 1.0000 37.000 1.4724 0.46050 0.45876 -0.3251 0.0022 1.0000