XFOIL Version 6.96 Calculated polar for: NLR-7301 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.500 -1.0979 0.10322 0.09898 -0.0445 1.0000 0.0356 -19.250 -1.1134 0.09689 0.09252 -0.0481 1.0000 0.0357 -19.000 -1.1241 0.09149 0.08700 -0.0510 1.0000 0.0359 -18.750 -1.1324 0.08659 0.08200 -0.0536 1.0000 0.0361 -18.500 -1.1395 0.08201 0.07732 -0.0560 1.0000 0.0363 -18.250 -1.1454 0.07770 0.07290 -0.0581 1.0000 0.0365 -18.000 -1.1502 0.07364 0.06873 -0.0601 1.0000 0.0367 -17.750 -1.1541 0.06982 0.06480 -0.0618 1.0000 0.0370 -17.500 -1.1568 0.06630 0.06117 -0.0632 1.0000 0.0373 -17.250 -1.1580 0.06302 0.05778 -0.0644 1.0000 0.0375 -17.000 -1.1586 0.05990 0.05455 -0.0654 1.0000 0.0378 -16.750 -1.1580 0.05702 0.05155 -0.0662 1.0000 0.0381 -16.500 -1.1564 0.05432 0.04872 -0.0669 1.0000 0.0384 -16.250 -1.1541 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0.1164 0.6762 2.750 0.5509 0.01364 0.00606 -0.0644 0.1087 0.6787 3.000 0.5772 0.01394 0.00629 -0.0641 0.1020 0.6813 3.250 0.6042 0.01417 0.00649 -0.0638 0.0962 0.6839 3.500 0.6306 0.01446 0.00672 -0.0635 0.0907 0.6863 3.750 0.6570 0.01476 0.00696 -0.0633 0.0855 0.6885 4.000 0.6836 0.01500 0.00717 -0.0630 0.0808 0.6907 4.250 0.7084 0.01532 0.00745 -0.0624 0.0769 0.6933 4.500 0.7343 0.01557 0.00772 -0.0620 0.0736 0.6957 4.750 0.7583 0.01599 0.00808 -0.0613 0.0709 0.6981 5.000 0.7835 0.01630 0.00842 -0.0607 0.0688 0.7008 5.250 0.8085 0.01663 0.00874 -0.0602 0.0669 0.7038 5.500 0.8327 0.01702 0.00910 -0.0595 0.0654 0.7067 5.750 0.8559 0.01752 0.00956 -0.0588 0.0640 0.7094 6.000 0.8807 0.01788 0.00994 -0.0582 0.0629 0.7117 6.250 0.9047 0.01822 0.01029 -0.0576 0.0618 0.7145 6.500 0.9279 0.01860 0.01068 -0.0568 0.0609 0.7172 6.750 0.9504 0.01903 0.01112 -0.0559 0.0601 0.7198 7.000 0.9716 0.01957 0.01165 -0.0548 0.0593 0.7226 7.250 0.9913 0.02023 0.01230 -0.0535 0.0586 0.7256 7.500 1.0128 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-0.0325 0.0511 0.7872 12.500 1.3938 0.03613 0.02922 -0.0312 0.0509 0.7906 12.750 1.4066 0.03728 0.03045 -0.0299 0.0506 0.7940 13.000 1.4183 0.03846 0.03171 -0.0285 0.0503 0.7971 13.250 1.4291 0.03968 0.03303 -0.0271 0.0501 0.8003 13.500 1.4389 0.04103 0.03447 -0.0256 0.0499 0.8037 13.750 1.4481 0.04247 0.03601 -0.0242 0.0498 0.8075 14.000 1.4563 0.04402 0.03765 -0.0229 0.0496 0.8113 14.250 1.4637 0.04568 0.03939 -0.0216 0.0495 0.8148 14.500 1.4696 0.04741 0.04124 -0.0203 0.0493 0.8183 14.750 1.4746 0.04929 0.04323 -0.0190 0.0492 0.8217 15.000 1.4780 0.05136 0.04540 -0.0178 0.0491 0.8253 15.250 1.4795 0.05367 0.04783 -0.0167 0.0490 0.8292 15.500 1.4788 0.05627 0.05055 -0.0157 0.0488 0.8328 15.750 1.4745 0.05917 0.05361 -0.0147 0.0488 0.8366 16.000 1.4666 0.06249 0.05711 -0.0139 0.0487 0.8403 16.250 1.4546 0.06636 0.06116 -0.0135 0.0486 0.8442 16.500 1.4380 0.07084 0.06584 -0.0135 0.0486 0.8478 16.750 1.4179 0.07593 0.07114 -0.0143 0.0486 0.8512 17.000 1.3941 0.08180 0.07725 -0.0160 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