XFOIL Version 6.94 Calculated polar for: Eif385mod3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1023 0.04360 0.03269 -0.0089 0.9993 0.5422 -2.750 -0.0681 0.04301 0.03210 -0.0109 0.9993 0.5407 -2.500 -0.0332 0.04253 0.03170 -0.0127 0.9993 0.5393 -2.250 0.0024 0.04217 0.03149 -0.0144 0.9993 0.5372 -2.000 0.0365 0.04195 0.03163 -0.0159 0.9993 0.5391 -1.500 0.2999 0.03540 0.02361 -0.0453 0.5222 0.5928 -1.250 0.3299 0.03561 0.02385 -0.0449 0.4989 0.6410 -1.000 0.3774 0.03578 0.02407 -0.0478 0.4803 1.0007 -0.750 0.4137 0.03696 0.02479 -0.0494 0.4680 1.0007 -0.500 0.4487 0.03827 0.02557 -0.0505 0.4579 1.0007 -0.250 0.4824 0.03949 0.02660 -0.0516 0.4488 1.0007 0.000 0.5154 0.04084 0.02766 -0.0524 0.4404 1.0007 0.250 0.5480 0.04230 0.02885 -0.0532 0.4341 1.0007 0.500 0.5793 0.04375 0.03034 -0.0542 0.4300 1.0007 0.750 0.6101 0.04536 0.03200 -0.0552 0.4280 1.0007 1.000 0.6396 0.04713 0.03384 -0.0562 0.4268 1.0007 1.250 0.6674 0.04909 0.03591 -0.0571 0.4263 1.0007 1.500 0.6945 0.05130 0.03825 -0.0582 0.4267 1.0007 1.750 0.7198 0.05377 0.04086 -0.0593 0.4277 1.0007 2.000 0.7430 0.05651 0.04373 -0.0604 0.4295 1.0007 2.250 0.7642 0.05952 0.04685 -0.0614 0.4318 1.0007 2.500 0.7842 0.06274 0.05015 -0.0624 0.4344 1.0007 2.750 0.7813 0.06774 0.05559 -0.0633 0.4410 1.0007 3.000 0.7646 0.07407 0.06220 -0.0638 0.4501 1.0007 3.250 0.7659 0.07900 0.06716 -0.0644 0.4563 1.0007 3.750 0.7087 0.09353 0.08187 -0.0655 0.4830 1.0007 4.000 0.6227 0.10445 0.09287 -0.0667 0.5203 1.0007 4.250 0.5865 0.11155 0.09999 -0.0688 0.5596 1.0007