XFOIL Version 6.96 Calculated polar for: E184 (8.33%) 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.5498 0.09894 0.09617 0.0242 1.0000 0.0493 -9.000 -0.5499 0.09433 0.09158 0.0224 1.0000 0.0508 -8.750 -0.6564 0.09274 0.08981 0.0191 1.0000 0.0468 -8.500 -0.6512 0.08935 0.08644 0.0186 1.0000 0.0482 -8.250 -0.6529 0.08502 0.08215 0.0155 1.0000 0.0491 -8.000 -0.6584 0.08034 0.07750 0.0117 1.0000 0.0501 -7.750 -0.6601 0.07525 0.07239 0.0078 1.0000 0.0512 -7.500 -0.6595 0.06976 0.06683 0.0038 1.0000 0.0531 -7.250 -0.6641 0.06431 0.06078 -0.0019 1.0000 0.0572 -7.000 -0.6613 0.05657 0.05297 -0.0033 1.0000 0.0586 -6.750 -0.6442 0.05321 0.04971 -0.0034 1.0000 0.0605 -6.500 -0.6273 0.05042 0.04686 -0.0038 1.0000 0.0647 -5.500 -0.5503 0.02779 0.02203 -0.0040 1.0000 0.0405 -5.250 -0.5175 0.02287 0.01617 -0.0043 0.9646 0.0351 -5.000 -0.4879 0.02087 0.01376 -0.0042 0.9412 0.0346 -4.750 -0.4660 0.01938 0.01208 -0.0028 0.9213 0.0359 -4.500 -0.4448 0.01845 0.01098 -0.0012 0.9052 0.0380 -4.250 -0.4226 0.01729 0.00965 0.0005 0.8915 0.0386 -4.000 -0.3998 0.01625 0.00849 0.0019 0.8792 0.0396 -3.750 -0.3769 0.01539 0.00753 0.0033 0.8681 0.0415 -3.500 -0.3540 0.01471 0.00672 0.0047 0.8581 0.0442 -3.250 -0.3317 0.01381 0.00577 0.0060 0.8482 0.0509 -3.000 -0.3093 0.01269 0.00493 0.0071 0.8387 0.1082 -2.750 -0.2899 0.01150 0.00453 0.0083 0.8304 0.2866 -2.500 -0.2704 0.01061 0.00430 0.0098 0.8215 0.4526 -2.250 -0.2543 0.00965 0.00419 0.0124 0.8134 0.6560 -2.000 -0.1795 0.00926 0.00450 0.0051 0.8081 0.9340 -1.750 -0.1236 0.00957 0.00459 0.0001 0.8016 0.9738 -1.500 -0.0619 0.00963 0.00441 -0.0068 0.7945 0.9962 -1.250 -0.0299 0.00958 0.00420 -0.0079 0.7868 1.0000 -1.000 -0.0065 0.00956 0.00405 -0.0072 0.7789 1.0000 -0.750 0.0179 0.00957 0.00394 -0.0067 0.7711 1.0000 -0.500 0.0418 0.00958 0.00384 -0.0060 0.7640 1.0000 -0.250 0.0671 0.00961 0.00380 -0.0057 0.7562 1.0000 0.000 0.0913 0.00965 0.00374 -0.0049 0.7495 1.0000 0.250 0.1171 0.00970 0.00375 -0.0047 0.7417 1.0000 0.500 0.1415 0.00976 0.00372 -0.0039 0.7353 1.0000 0.750 0.1676 0.00983 0.00377 -0.0037 0.7273 1.0000 1.000 0.1921 0.00990 0.00377 -0.0030 0.7210 1.0000 1.250 0.2182 0.00997 0.00384 -0.0027 0.7123 1.0000 1.500 0.2429 0.01004 0.00387 -0.0019 0.7049 1.0000 1.750 0.2685 0.01010 0.00392 -0.0015 0.6957 1.0000 2.000 0.2939 0.01016 0.00398 -0.0009 0.6867 1.0000 2.250 0.3184 0.01021 0.00398 0.0000 0.6783 1.0000 2.500 0.3443 0.01025 0.00406 0.0005 0.6675 1.0000 2.750 0.3698 0.01029 0.00413 0.0011 0.6571 1.0000 3.000 0.3950 0.01031 0.00416 0.0018 0.6464 1.0000 3.250 0.4200 0.01033 0.00416 0.0026 0.6356 1.0000 3.500 0.4456 0.01033 0.00423 0.0033 0.6229 1.0000 3.750 0.4711 0.01032 0.00426 0.0040 0.6090 1.0000 4.000 0.4965 0.01030 0.00428 0.0047 0.5938 1.0000 4.250 0.5225 0.01028 0.00435 0.0053 0.5748 1.0000 4.500 0.5480 0.01026 0.00440 0.0060 0.5536 1.0000 4.750 0.5737 0.01027 0.00448 0.0066 0.5268 1.0000 5.000 0.5990 0.01033 0.00450 0.0073 0.4782 1.0000 5.250 0.6224 0.01096 0.00458 0.0078 0.3480 1.0000 5.500 0.6445 0.01242 0.00529 0.0076 0.2045 1.0000 5.750 0.6643 0.01456 0.00656 0.0074 0.0592 1.0000 6.000 0.6856 0.01626 0.00815 0.0081 0.0320 1.0000 6.250 0.7077 0.01736 0.00934 0.0088 0.0276 1.0000 6.500 0.7271 0.01890 0.01092 0.0098 0.0252 1.0000 6.750 0.7447 0.02106 0.01312 0.0112 0.0241 1.0000 7.000 0.7659 0.02267 0.01484 0.0125 0.0238 1.0000 7.250 0.7872 0.02467 0.01703 0.0137 0.0236 1.0000 7.500 0.8085 0.02668 0.01927 0.0149 0.0233 1.0000 7.750 0.8294 0.02824 0.02110 0.0159 0.0220 1.0000 8.000 0.8481 0.03064 0.02383 0.0171 0.0215 1.0000 8.250 0.8634 0.03399 0.02760 0.0185 0.0219 1.0000 8.500 0.8741 0.03799 0.03200 0.0198 0.0228 1.0000 8.750 0.8847 0.04283 0.03705 0.0208 0.0246 1.0000 16.250 0.6077 0.18140 0.17801 -0.0282 0.0340 1.0000 16.500 0.6093 0.18445 0.18106 -0.0302 0.0339 1.0000