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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1078 0.03357 0.02260 -0.0425 0.4567 0.4876 -2.750 0.1364 0.03440 0.02231 -0.0427 0.4076 0.4771 -2.500 0.1673 0.03472 0.02206 -0.0431 0.3877 0.4708 -2.250 0.2000 0.03496 0.02188 -0.0437 0.3741 0.4663 -2.000 0.2336 0.03536 0.02185 -0.0444 0.3623 0.4611 -1.750 0.2681 0.03586 0.02195 -0.0452 0.3532 0.4563 -1.500 0.3027 0.03603 0.02203 -0.0459 0.3459 0.4563 -1.250 0.3378 0.03630 0.02222 -0.0467 0.3410 0.4572 -1.000 0.3725 0.03655 0.02250 -0.0475 0.3374 0.4604 -0.750 0.4072 0.03693 0.02293 -0.0483 0.3345 0.4645 -0.500 0.4428 0.03748 0.02354 -0.0494 0.3323 0.4677 -0.250 0.4769 0.03808 0.02431 -0.0503 0.3307 0.4769 0.000 0.5093 0.03873 0.02527 -0.0510 0.3296 0.4923 0.500 0.5797 0.03971 0.02748 -0.0529 0.3290 1.0007 0.750 0.6104 0.04099 0.02877 -0.0534 0.3295 1.0007 1.000 0.6403 0.04236 0.03017 -0.0538 0.3300 1.0007 1.250 0.6693 0.04388 0.03175 -0.0542 0.3301 1.0007 1.500 0.6974 0.04556 0.03347 -0.0546 0.3300 1.0007 1.750 0.7247 0.04738 0.03534 -0.0549 0.3298 1.0007 2.000 0.7511 0.04922 0.03728 -0.0553 0.3299 1.0007 2.250 0.7764 0.05100 0.03927 -0.0556 0.3315 1.0007 2.500 0.7982 0.05318 0.04188 -0.0560 0.3364 1.0007 2.750 0.8179 0.05601 0.04503 -0.0565 0.3420 1.0007 3.000 0.8370 0.05910 0.04831 -0.0569 0.3475 1.0007 3.250 0.8562 0.06239 0.05168 -0.0572 0.3525 1.0007 3.500 0.8800 0.06591 0.05513 -0.0578 0.3564 1.0007 3.750 0.8531 0.07262 0.06277 -0.0585 0.3778 1.0007 4.000 0.8832 0.07626 0.06627 -0.0596 0.3841 1.0007 5.000 0.4853 0.12568 0.11690 -0.0747 0.7530 1.0007