XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0139 0.02288 0.01275 -0.0253 0.9994 1.0006 -2.750 0.0076 0.02346 0.01313 -0.0251 0.9994 1.0006 -2.500 0.0291 0.02410 0.01358 -0.0248 0.9994 1.0006 -2.250 0.0507 0.02479 0.01412 -0.0247 0.9994 1.0006 -2.000 0.0724 0.02553 0.01472 -0.0245 0.9994 1.0006 -1.750 0.0941 0.02630 0.01537 -0.0245 0.9994 1.0006 -1.500 0.1156 0.02711 0.01608 -0.0244 0.9994 1.0006 -1.250 0.1371 0.02795 0.01685 -0.0244 0.9994 1.0006 -1.000 0.1584 0.02884 0.01768 -0.0245 0.9994 1.0006 -0.750 0.1796 0.02976 0.01856 -0.0246 0.9994 1.0006 -0.500 0.2005 0.03073 0.01951 -0.0247 0.9994 1.0006 -0.250 0.2212 0.03175 0.02052 -0.0249 0.9994 1.0006 0.000 0.2416 0.03283 0.02162 -0.0251 0.9994 1.0006 0.250 0.2616 0.03399 0.02282 -0.0254 0.9994 1.0006 0.500 0.3334 0.03380 0.02271 -0.0350 0.9800 1.0006 0.750 0.4308 0.03211 0.02120 -0.0480 0.9442 1.0006 1.000 0.5354 0.02933 0.01870 -0.0599 0.9013 1.0006 1.250 0.5808 0.02812 0.01763 -0.0591 0.8464 1.0006 1.500 0.6011 0.02745 0.01696 -0.0528 0.7807 1.0006 1.750 0.6144 0.02688 0.01621 -0.0454 0.6930 1.0006 2.000 0.6317 0.02680 0.01560 -0.0399 0.5995 1.0006 2.250 0.6542 0.02738 0.01559 -0.0374 0.5280 1.0006 2.500 0.6797 0.02829 0.01603 -0.0362 0.4796 1.0006 2.750 0.7067 0.02939 0.01686 -0.0357 0.4442 1.0006 3.000 0.7342 0.03060 0.01775 -0.0353 0.4158 1.0006 3.250 0.7623 0.03195 0.01901 -0.0354 0.3908 1.0006 3.500 0.7903 0.03343 0.02034 -0.0354 0.3703 1.0006 3.750 0.8182 0.03506 0.02185 -0.0354 0.3520 1.0006 4.000 0.8468 0.03687 0.02374 -0.0358 0.3371 1.0006 4.250 0.8754 0.03887 0.02593 -0.0364 0.3246 1.0006 4.500 0.9034 0.04100 0.02815 -0.0368 0.3136 1.0006 4.750 0.9305 0.04300 0.03024 -0.0371 0.3014 1.0006 5.000 0.9564 0.04519 0.03279 -0.0376 0.2893 1.0006