XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY MS(1)-0317 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 ------ -------- --------- --------- -------- -------- -------- -14.750 -0.5797 0.08953 0.08539 -0.0633 1.0000 0.0516 -14.500 -0.5697 0.08797 0.08386 -0.0634 1.0000 0.0523 -14.250 -0.9042 0.05918 0.05355 -0.0733 1.0000 0.0459 -14.000 -0.9118 0.05595 0.05018 -0.0734 1.0000 0.0462 -13.750 -0.9170 0.05308 0.04717 -0.0731 1.0000 0.0465 -13.500 -0.9206 0.05054 0.04451 -0.0726 1.0000 0.0470 -13.250 -0.9253 0.04809 0.04191 -0.0717 1.0000 0.0475 -13.000 -0.9312 0.04573 0.03937 -0.0703 1.0000 0.0481 -12.750 -0.9387 0.04358 0.03702 -0.0681 1.0000 0.0487 -12.500 -0.9491 0.04179 0.03505 -0.0648 1.0000 0.0492 -12.250 -0.9589 0.04020 0.03326 -0.0611 1.0000 0.0497 -12.000 -0.9697 0.03894 0.03180 -0.0566 1.0000 0.0501 -11.750 -0.9827 0.03800 0.03066 -0.0515 1.0000 0.0505 -11.500 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0.6273 0.01659 0.01061 -0.0610 0.6811 0.8021 3.000 0.6535 0.01639 0.01041 -0.0603 0.6670 0.8047 3.250 0.6806 0.01623 0.01025 -0.0601 0.6496 0.8074 3.500 0.7090 0.01607 0.01006 -0.0602 0.6288 0.8105 3.750 0.7379 0.01599 0.00993 -0.0606 0.6004 0.8142 4.000 0.7661 0.01598 0.00978 -0.0608 0.5600 0.8173 4.250 0.7847 0.01617 0.00968 -0.0589 0.4975 0.8195 4.500 0.7960 0.01689 0.00989 -0.0559 0.4117 0.8223 4.750 0.8066 0.01783 0.01039 -0.0533 0.3383 0.8255 5.000 0.8209 0.01872 0.01092 -0.0516 0.2852 0.8287 5.250 0.8393 0.01950 0.01146 -0.0508 0.2505 0.8317 5.500 0.8604 0.02026 0.01200 -0.0505 0.2271 0.8346 5.750 0.8761 0.02076 0.01243 -0.0486 0.2122 0.8370 6.000 0.8927 0.02131 0.01293 -0.0469 0.2009 0.8399 6.250 0.9097 0.02195 0.01346 -0.0455 0.1917 0.8429 6.500 0.9305 0.02246 0.01398 -0.0447 0.1840 0.8459 6.750 0.9509 0.02304 0.01450 -0.0441 0.1774 0.8489 7.000 0.9738 0.02376 0.01515 -0.0440 0.1716 0.8517 7.250 0.9917 0.02416 0.01560 -0.0426 0.1668 0.8542 7.500 1.0102 0.02469 0.01612 -0.0413 0.1627 0.8572 7.750 1.0340 0.02556 0.01685 -0.0411 0.1584 0.8605 8.000 1.0568 0.02606 0.01745 -0.0407 0.1556 0.8639 8.250 1.0820 0.02666 0.01810 -0.0408 0.1524 0.8670 8.500 1.1054 0.02725 0.01871 -0.0406 0.1493 0.8698 8.750 1.1270 0.02785 0.01929 -0.0399 0.1466 0.8729 9.000 1.1578 0.02881 0.02016 -0.0407 0.1435 0.8761 9.250 1.1795 0.02941 0.02089 -0.0402 0.1417 0.8801 9.500 1.2041 0.03013 0.02170 -0.0402 0.1396 0.8837 9.750 1.2257 0.03078 0.02243 -0.0396 0.1374 0.8869 10.000 1.2465 0.03143 0.02313 -0.0389 0.1352 0.8905 10.250 1.2705 0.03216 0.02386 -0.0388 0.1332 0.8944 10.500 1.3035 0.03324 0.02489 -0.0402 0.1310 0.8983 10.750 1.3291 0.03438 0.02613 -0.0404 0.1291 0.9024 11.000 1.3424 0.03509 0.02700 -0.0386 0.1277 0.9075 11.250 1.3599 0.03597 0.02803 -0.0376 0.1259 0.9126 11.500 1.3765 0.03685 0.02904 -0.0365 0.1241 0.9176 11.750 1.3918 0.03766 0.02994 -0.0352 0.1223 0.9234 12.000 1.4112 0.03853 0.03083 -0.0347 0.1204 0.9295 12.250 1.4341 0.03945 0.03172 -0.0345 0.1186 0.9359 12.500 1.4545 0.04105 0.03342 -0.0342 0.1167 0.9435 12.750 1.4558 0.04196 0.03456 -0.0311 0.1155 0.9557 13.000 1.4589 0.04301 0.03582 -0.0286 0.1141 0.9796 13.250 1.4705 0.04441 0.03739 -0.0280 0.1123 1.0000 13.500 1.4837 0.04586 0.03895 -0.0278 0.1106 1.0000 13.750 1.4981 0.04723 0.04040 -0.0276 0.1089 1.0000 14.000 1.5208 0.04844 0.04155 -0.0283 0.1072 1.0000 14.250 1.5417 0.05073 0.04389 -0.0289 0.1054 1.0000 14.500 1.5340 0.05302 0.04647 -0.0270 0.1044 1.0000 14.750 1.5262 0.05564 0.04936 -0.0255 0.1032 1.0000 15.000 1.5198 0.05838 0.05232 -0.0243 0.1018 1.0000 15.250 1.5169 0.06097 0.05508 -0.0236 0.1004 1.0000 15.500 1.5188 0.06321 0.05742 -0.0231 0.0990 1.0000 15.750 1.5223 0.06531 0.05958 -0.0228 0.0977 1.0000 16.000 1.5357 0.06678 0.06102 -0.0228 0.0964 1.0000 16.250 1.5442 0.06957 0.06383 -0.0227 0.0949 1.0000 16.500 1.5114 0.07475 0.06935 -0.0221 0.0943 1.0000 16.750 1.4734 0.08112 0.07605 -0.0226 0.0938 1.0000 17.000 1.4267 0.08938 0.08467 -0.0246 0.0933 1.0000 17.250 1.3572 0.10216 0.09787 -0.0299 0.0930 1.0000