XFOIL Version 6.94 Calculated polar for: NLR-7301 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 0.2871 0.00969 0.00509 -0.0726 0.7614 0.6580 0.500 0.3433 0.00944 0.00476 -0.0718 0.7173 0.6629 1.000 0.3871 0.01018 0.00472 -0.0687 0.5378 0.6674 1.500 0.4185 0.01234 0.00552 -0.0648 0.2514 0.6718 2.000 0.4662 0.01340 0.00600 -0.0636 0.1480 0.6767 2.500 0.5196 0.01402 0.00639 -0.0631 0.1200 0.6814 3.000 0.5724 0.01448 0.00672 -0.0626 0.1042 0.6860 3.500 0.6237 0.01507 0.00724 -0.0616 0.0921 0.6903 4.000 0.6747 0.01571 0.00782 -0.0605 0.0824 0.6951 5.000 0.7763 0.01709 0.00911 -0.0586 0.0703 0.7059 5.500 0.8233 0.01797 0.00994 -0.0571 0.0669 0.7111 6.000 0.8711 0.01871 0.01075 -0.0556 0.0645 0.7161 6.500 0.9174 0.01958 0.01162 -0.0539 0.0625 0.7217 7.000 0.9598 0.02090 0.01288 -0.0518 0.0605 0.7280 7.500 1.0042 0.02177 0.01381 -0.0500 0.0593 0.7337 8.000 1.0474 0.02282 0.01497 -0.0479 0.0581 0.7394 8.500 1.0906 0.02402 0.01624 -0.0460 0.0570 0.7454 9.000 1.1337 0.02531 0.01756 -0.0443 0.0561 0.7518 9.500 1.1765 0.02674 0.01900 -0.0427 0.0553 0.7575 10.000 1.2208 0.02864 0.02096 -0.0413 0.0545 0.7633 10.500 1.2628 0.03071 0.02318 -0.0396 0.0540 0.7699 11.000 1.2985 0.03262 0.02527 -0.0372 0.0536 0.7769 11.500 1.3305 0.03478 0.02766 -0.0346 0.0532 0.7830 12.000 1.3565 0.03729 0.03044 -0.0314 0.0528 0.7893 12.500 1.3760 0.04006 0.03347 -0.0280 0.0522 0.7962 13.000 1.3898 0.04329 0.03696 -0.0246 0.0517 0.8025 13.500 1.3946 0.04724 0.04125 -0.0210 0.0514 0.8090 14.000 1.3893 0.05219 0.04655 -0.0177 0.0511 0.8162 14.500 1.3702 0.05874 0.05350 -0.0151 0.0510 0.8226 15.000 1.3292 0.06806 0.06332 -0.0142 0.0510 0.8284 15.500 1.1868 0.09418 0.09039 -0.0249 0.0516 0.8301