XFOIL Version 6.94 Calculated polar for: Eif385mod4 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0951 0.03568 0.02490 -0.0372 0.3725 0.5384 -2.750 0.1216 0.03509 0.02398 -0.0372 0.3581 0.5531 -2.500 0.1517 0.03461 0.02324 -0.0377 0.3474 0.5678 -2.250 0.1827 0.03412 0.02262 -0.0381 0.3374 0.5800 -1.750 0.2521 0.03399 0.02214 -0.0404 0.3226 0.5985 -1.500 0.2886 0.03389 0.02206 -0.0417 0.3186 0.6110 -1.250 0.3255 0.03386 0.02215 -0.0431 0.3155 0.6247 -1.000 0.3639 0.03403 0.02242 -0.0448 0.3128 0.6334 -0.750 0.4006 0.03425 0.02284 -0.0461 0.3109 0.6490 -0.500 0.4353 0.03450 0.02339 -0.0468 0.3094 0.6880 -0.250 0.4832 0.03471 0.02400 -0.0502 0.3083 1.0007 0.000 0.5183 0.03579 0.02494 -0.0515 0.3078 1.0007 0.250 0.5519 0.03695 0.02601 -0.0525 0.3079 1.0007 0.500 0.5843 0.03817 0.02721 -0.0533 0.3085 1.0007 0.750 0.6155 0.03947 0.02853 -0.0540 0.3091 1.0007 1.000 0.6458 0.04086 0.02995 -0.0546 0.3094 1.0007 1.250 0.6710 0.04245 0.03153 -0.0542 0.3094 0.8031 1.500 0.6985 0.04476 0.03324 -0.0549 0.3092 0.5678 1.750 0.7283 0.04672 0.03508 -0.0557 0.3095 0.5505 2.000 0.7562 0.04874 0.03713 -0.0563 0.3110 0.5389 2.250 0.7823 0.05089 0.03945 -0.0568 0.3145 0.5334 2.500 0.8070 0.05318 0.04196 -0.0572 0.3182 0.5323 2.750 0.8310 0.05573 0.04466 -0.0576 0.3219 0.5317 3.000 0.8553 0.05875 0.04769 -0.0581 0.3253 0.5276 3.250 0.8685 0.06140 0.05107 -0.0587 0.3370 0.5228 3.500 0.8804 0.06577 0.05569 -0.0594 0.3478 0.5180 3.750 0.8906 0.07007 0.06029 -0.0604 0.3609 0.5138 4.000 0.8903 0.07697 0.06744 -0.0620 0.3815 0.5114 4.500 0.5117 0.12201 0.11422 -0.0802 0.7330 0.5674 4.750 0.5251 0.12475 0.11655 -0.0802 0.7163 0.5465 5.000 0.5383 0.12736 0.11895 -0.0799 0.6994 0.5374