XFOIL Version 6.94 Calculated polar for: ARA-D BL 13% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2118 0.04887 0.04098 -0.0119 1.0000 0.3365 -2.750 -0.1919 0.04682 0.03899 -0.0117 1.0000 0.3437 -2.500 -0.1675 0.04508 0.03711 -0.0127 1.0000 0.3524 -2.250 -0.1402 0.04353 0.03528 -0.0144 1.0000 0.3630 -2.000 -0.1174 0.04233 0.03416 -0.0143 1.0000 0.3767 -1.750 -0.0903 0.04135 0.03312 -0.0154 1.0000 0.3930 -1.500 -0.0643 0.04085 0.03262 -0.0165 1.0000 0.4132 -1.250 0.1858 0.03197 0.02403 -0.0299 0.5392 0.6222 -1.000 0.2546 0.03083 0.02257 -0.0332 0.4692 1.0000 -0.750 0.2869 0.03208 0.02268 -0.0333 0.4430 1.0000 -0.500 0.3165 0.03338 0.02335 -0.0331 0.4242 1.0000 -0.250 0.3451 0.03470 0.02418 -0.0327 0.4090 1.0000 0.000 0.3732 0.03605 0.02513 -0.0323 0.3951 1.0000 0.250 0.4011 0.03762 0.02653 -0.0323 0.3843 1.0000 0.500 0.4289 0.03928 0.02805 -0.0324 0.3763 1.0000 0.750 0.4575 0.04111 0.02966 -0.0325 0.3710 1.0000 1.000 0.4842 0.04339 0.03198 -0.0331 0.3678 1.0000 1.250 0.5089 0.04602 0.03474 -0.0341 0.3657 1.0000 1.500 0.5311 0.04894 0.03779 -0.0350 0.3637 1.0000 1.750 0.5502 0.05216 0.04113 -0.0360 0.3617 1.0000 2.000 0.5655 0.05577 0.04485 -0.0372 0.3604 1.0000 2.250 0.5761 0.05995 0.04915 -0.0386 0.3610 1.0000 2.500 0.5826 0.06465 0.05394 -0.0403 0.3641 1.0000 2.750 0.5934 0.06904 0.05831 -0.0415 0.3682 1.0000 3.000 0.5544 0.07850 0.06811 -0.0478 0.3869 1.0000 3.500 0.4725 0.09527 0.08505 -0.0596 0.4519 1.0000 4.500 0.3443 0.11538 0.10539 -0.0772 0.7772 1.0000 4.750 0.3434 0.11665 0.10655 -0.0755 0.7650 1.0000 5.000 0.3704 0.12157 0.11128 -0.0778 0.7527 1.0000