XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY NLF(2)-0415 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 ------ -------- --------- --------- -------- -------- -------- -13.000 -0.1796 0.09170 0.08766 -0.1260 0.9207 0.0244 -12.750 -0.1969 0.07986 0.07579 -0.1308 0.9201 0.0177 -12.500 -0.2028 0.07402 0.06993 -0.1336 0.9190 0.0174 -12.250 -0.2956 0.07944 0.07521 -0.1363 0.9292 0.0173 -12.000 -0.3052 0.07361 0.06930 -0.1405 0.9260 0.0171 -11.750 -0.3177 0.06805 0.06362 -0.1442 0.9228 0.0169 -11.500 -0.3303 0.06299 0.05841 -0.1474 0.9197 0.0167 -11.250 -0.3428 0.05896 0.05424 -0.1490 0.9161 0.0165 -11.000 -0.3584 0.05536 0.05049 -0.1490 0.9119 0.0163 -10.750 -0.3724 0.05199 0.04693 -0.1486 0.9083 0.0162 -10.500 -0.3827 0.04897 0.04371 -0.1479 0.9052 0.0160 -10.250 -0.3945 0.04629 0.04082 -0.1459 0.9020 0.0160 -10.000 -0.4075 0.04428 0.03864 -0.1424 0.8983 0.0159 -9.750 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0.6406 -0.250 0.1225 0.02223 0.01495 -0.0689 0.8321 0.6413 0.000 0.1515 0.02229 0.01498 -0.0696 0.8313 0.6420 0.250 0.1816 0.02234 0.01501 -0.0704 0.8307 0.6427 0.500 0.2119 0.02240 0.01505 -0.0713 0.8301 0.6435 1.000 0.2279 0.02340 0.01605 -0.0653 0.8208 0.6458 1.250 0.2554 0.02351 0.01616 -0.0657 0.8193 0.6469 1.500 0.2841 0.02361 0.01627 -0.0663 0.8181 0.6478 1.750 0.3137 0.02370 0.01637 -0.0671 0.8171 0.6487 2.000 0.3444 0.02377 0.01646 -0.0680 0.8163 0.6495 2.250 0.3757 0.02385 0.01657 -0.0691 0.8155 0.6504 2.750 0.3994 0.02468 0.01751 -0.0644 0.8050 0.6520 3.000 0.4292 0.02473 0.01764 -0.0651 0.8035 0.6528 3.250 0.4852 0.02359 0.01660 -0.0695 0.8013 0.6536 3.750 0.5354 0.02233 0.01548 -0.0677 0.7801 0.6554 4.000 0.5841 0.02089 0.01416 -0.0706 0.7740 0.6565 5.000 0.7184 0.01573 0.00809 -0.0696 0.4830 0.6619 5.250 0.6992 0.01754 0.00917 -0.0618 0.3527 0.6633 5.500 0.6970 0.01891 0.01004 -0.0574 0.2588 0.6645 6.000 0.7195 0.02064 0.01129 -0.0530 0.1576 0.6669 6.250 0.7358 0.02127 0.01185 -0.0516 0.1310 0.6680 6.500 0.7532 0.02187 0.01239 -0.0504 0.1120 0.6693 6.750 0.7713 0.02244 0.01294 -0.0492 0.0981 0.6706 7.000 0.7898 0.02301 0.01351 -0.0482 0.0874 0.6721 7.250 0.8087 0.02356 0.01413 -0.0472 0.0793 0.6737 7.500 0.8267 0.02422 0.01477 -0.0461 0.0719 0.6753 7.750 0.8463 0.02476 0.01540 -0.0453 0.0658 0.6771 8.000 0.8644 0.02546 0.01611 -0.0442 0.0609 0.6789 8.500 0.9024 0.02670 0.01754 -0.0424 0.0523 0.6824 8.750 0.9200 0.02746 0.01831 -0.0413 0.0477 0.6842 9.000 0.9406 0.02793 0.01891 -0.0407 0.0426 0.6865 9.250 0.9588 0.02864 0.01962 -0.0399 0.0379 0.6888 9.500 0.9790 0.02920 0.02033 -0.0392 0.0336 0.6915 9.750 0.9973 0.02996 0.02112 -0.0384 0.0299 0.6941 10.000 1.0160 0.03072 0.02202 -0.0375 0.0266 0.6965 10.250 1.0336 0.03154 0.02292 -0.0366 0.0234 0.6988 10.500 1.0503 0.03257 0.02408 -0.0355 0.0208 0.7014 10.750 1.0670 0.03357 0.02521 -0.0345 0.0186 0.7043 11.000 1.0820 0.03470 0.02642 -0.0334 0.0173 0.7075 11.500 1.1130 0.03709 0.02917 -0.0312 0.0147 0.7139 11.750 1.1270 0.03825 0.03046 -0.0301 0.0138 0.7175 12.000 1.1387 0.03964 0.03197 -0.0288 0.0131 0.7216 12.500 1.1598 0.04317 0.03595 -0.0260 0.0121 0.7299 12.750 1.1682 0.04513 0.03818 -0.0245 0.0117 0.7349 13.250 1.1788 0.04952 0.04310 -0.0212 0.0111 0.7463 13.500 1.1804 0.05199 0.04582 -0.0195 0.0109 0.7532 13.750 1.1795 0.05469 0.04879 -0.0178 0.0107 0.7605 14.000 1.1763 0.05760 0.05197 -0.0162 0.0105 0.7691 14.250 1.1705 0.06078 0.05542 -0.0147 0.0104 0.7789 14.500 1.1617 0.06440 0.05933 -0.0134 0.0103 0.7904 14.750 1.1509 0.06830 0.06351 -0.0125 0.0102 0.8053 15.000 1.1372 0.07266 0.06817 -0.0119 0.0101 0.8268 15.250 1.1187 0.07720 0.07313 -0.0111 0.0101 0.9015 15.500 1.0959 0.08342 0.07963 -0.0123 0.0101 1.0000 15.750 1.0714 0.09120 0.08768 -0.0153 0.0102 1.0000 16.000 1.0396 0.10167 0.09843 -0.0209 0.0102 1.0000