XFOIL Version 6.96 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 = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.500 -0.5512 0.08856 0.08310 -0.0462 1.0000 0.0332 -11.250 -0.5908 0.07501 0.06947 -0.0563 1.0000 0.0325 -11.000 -0.6254 0.06676 0.06105 -0.0617 1.0000 0.0321 -10.750 -0.6540 0.06127 0.05540 -0.0636 1.0000 0.0319 -10.500 -0.6806 0.05753 0.05151 -0.0626 1.0000 0.0318 -10.250 -0.7025 0.05438 0.04820 -0.0603 1.0000 0.0318 -10.000 -0.7194 0.05154 0.04515 -0.0574 1.0000 0.0319 -9.750 -0.7321 0.04890 0.04228 -0.0542 1.0000 0.0320 -9.500 -0.7402 0.04640 0.03951 -0.0511 1.0000 0.0325 -9.250 -0.7440 0.04388 0.03665 -0.0484 1.0000 0.0332 -9.000 -0.7412 0.04099 0.03327 -0.0466 0.9993 0.0342 -8.750 -0.7172 0.03818 0.03007 -0.0483 0.9952 0.0353 -8.500 -0.6904 0.03628 0.02800 -0.0497 0.9914 0.0363 -8.250 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0.01786 0.01006 -0.0629 0.7944 0.8205 1.500 0.4633 0.01762 0.00983 -0.0627 0.7802 0.8254 1.750 0.4885 0.01740 0.00963 -0.0616 0.7655 0.8291 2.000 0.5148 0.01719 0.00945 -0.0608 0.7493 0.8336 2.250 0.5426 0.01701 0.00928 -0.0605 0.7303 0.8386 2.500 0.5696 0.01681 0.00907 -0.0598 0.7087 0.8428 2.750 0.5932 0.01669 0.00898 -0.0586 0.6804 0.8472 3.000 0.6192 0.01657 0.00880 -0.0578 0.6399 0.8522 3.250 0.6467 0.01651 0.00846 -0.0571 0.5703 0.8570 3.500 0.6661 0.01688 0.00826 -0.0549 0.4756 0.8612 3.750 0.6835 0.01763 0.00851 -0.0531 0.3936 0.8665 4.250 0.7219 0.01900 0.00927 -0.0506 0.2871 0.8769 4.500 0.7423 0.01960 0.00967 -0.0496 0.2533 0.8826 4.750 0.7645 0.02015 0.01010 -0.0490 0.2280 0.8885 5.000 0.7846 0.02064 0.01049 -0.0478 0.2101 0.8943 5.250 0.8068 0.02115 0.01095 -0.0471 0.1965 0.9006 5.500 0.8273 0.02165 0.01140 -0.0460 0.1862 0.9071 5.750 0.8485 0.02214 0.01189 -0.0451 0.1768 0.9147 6.000 0.8691 0.02263 0.01237 -0.0440 0.1691 0.9225 6.250 0.8906 0.02313 0.01288 -0.0432 0.1619 0.9313 6.500 0.9106 0.02367 0.01341 -0.0421 0.1564 0.9419 6.750 0.9336 0.02417 0.01398 -0.0416 0.1512 0.9555 7.000 0.9578 0.02474 0.01458 -0.0414 0.1466 0.9808 7.250 0.9838 0.02555 0.01533 -0.0419 0.1425 1.0000 7.500 1.0114 0.02629 0.01617 -0.0425 0.1381 1.0000 7.750 1.0381 0.02707 0.01700 -0.0429 0.1339 1.0000 8.000 1.0642 0.02792 0.01785 -0.0433 0.1304 1.0000 8.250 1.0913 0.02894 0.01882 -0.0439 0.1277 1.0000 8.500 1.1180 0.02984 0.01989 -0.0442 0.1248 1.0000 8.750 1.1436 0.03078 0.02095 -0.0445 0.1216 1.0000 9.000 1.1683 0.03172 0.02196 -0.0446 0.1186 1.0000 9.250 1.1927 0.03270 0.02294 -0.0447 0.1160 1.0000 9.500 1.2187 0.03393 0.02416 -0.0451 0.1137 1.0000 9.750 1.2407 0.03506 0.02557 -0.0447 0.1113 1.0000 10.000 1.2616 0.03622 0.02694 -0.0442 0.1085 1.0000 10.250 1.2814 0.03733 0.02818 -0.0437 0.1057 1.0000 10.500 1.3012 0.03841 0.02932 -0.0432 0.1034 1.0000 10.750 1.3227 0.03958 0.03047 -0.0430 0.1011 1.0000 11.000 1.3366 0.04103 0.03219 -0.0417 0.0987 1.0000 11.250 1.3479 0.04251 0.03397 -0.0401 0.0962 1.0000 11.500 1.3573 0.04387 0.03554 -0.0382 0.0939 1.0000 11.750 1.3673 0.04511 0.03690 -0.0366 0.0917 1.0000 12.000 1.3795 0.04629 0.03814 -0.0353 0.0898 1.0000 12.250 1.3943 0.04769 0.03952 -0.0344 0.0880 1.0000 12.500 1.3900 0.04990 0.04214 -0.0316 0.0862 1.0000 12.750 1.3859 0.05226 0.04483 -0.0293 0.0842 1.0000 13.000 1.3829 0.05460 0.04744 -0.0275 0.0824 1.0000 13.250 1.3811 0.05687 0.04990 -0.0259 0.0807 1.0000 13.500 1.3820 0.05892 0.05207 -0.0248 0.0791 1.0000 13.750 1.3885 0.06055 0.05374 -0.0239 0.0776 1.0000 14.000 1.3896 0.06289 0.05615 -0.0231 0.0762 1.0000 14.250 1.3656 0.06774 0.06141 -0.0223 0.0752 1.0000 14.500 1.3383 0.07343 0.06747 -0.0225 0.0742 1.0000 14.750 1.3059 0.08036 0.07474 -0.0240 0.0734 1.0000 15.000 1.2620 0.08981 0.08452 -0.0278 0.0731 1.0000