XFOIL Version 6.96 Calculated polar for: NACA 66-206 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.5083 0.08313 0.08149 -0.0211 1.0000 0.0039 -7.750 -0.5108 0.07951 0.07790 -0.0236 1.0000 0.0039 -7.500 -0.5117 0.07533 0.07372 -0.0275 1.0000 0.0039 -7.250 -0.5081 0.07117 0.06953 -0.0302 1.0000 0.0039 -7.000 -0.5040 0.06695 0.06528 -0.0322 1.0000 0.0039 -6.750 -0.4987 0.06298 0.06127 -0.0333 1.0000 0.0039 -6.500 -0.4790 0.05756 0.05574 -0.0379 0.9928 0.0039 -6.250 -0.4562 0.05228 0.05033 -0.0420 0.9878 0.0039 -6.000 -0.4323 0.04711 0.04501 -0.0454 0.9823 0.0039 -5.750 -0.4066 0.04219 0.03990 -0.0483 0.9768 0.0039 -5.500 -0.3887 0.03462 0.03198 -0.0510 0.9673 0.0033 -5.250 -0.3657 0.02994 0.02700 -0.0515 0.9587 0.0028 -5.000 -0.3426 0.02558 0.02229 -0.0510 0.9506 0.0025 -4.750 -0.3199 0.02067 0.01692 -0.0496 0.9420 0.0021 -4.500 -0.2988 0.01164 0.00682 -0.0462 0.9338 0.0016 -4.250 -0.2744 0.00989 0.00475 -0.0453 0.9277 0.0015 -4.000 -0.2491 0.00896 0.00363 -0.0448 0.9215 0.0015 -3.750 -0.2233 0.00833 0.00284 -0.0444 0.9158 0.0015 -3.500 -0.1968 0.00788 0.00225 -0.0441 0.9104 0.0017 -3.250 -0.1699 0.00760 0.00186 -0.0439 0.9047 0.0021 -3.000 -0.1430 0.00739 0.00165 -0.0438 0.8996 0.0030 -2.750 -0.1158 0.00724 0.00141 -0.0438 0.8942 0.0029 -2.500 -0.0886 0.00711 0.00125 -0.0437 0.8894 0.0036 -2.250 -0.0614 0.00696 0.00112 -0.0438 0.8845 0.0080 -2.000 -0.0341 0.00692 0.00115 -0.0439 0.8793 0.0111 -1.750 -0.0069 0.00674 0.00091 -0.0438 0.8746 0.0168 -1.500 0.0207 0.00664 0.00083 -0.0438 0.8696 0.0204 -1.250 0.0481 0.00654 0.00075 -0.0438 0.8649 0.0294 -1.000 0.0753 0.00634 0.00067 -0.0439 0.8604 0.0776 -0.750 0.1008 0.00549 0.00057 -0.0441 0.8549 0.3373 -0.500 0.1225 0.00416 0.00054 -0.0436 0.8489 0.7538 -0.250 0.1481 0.00404 0.00059 -0.0431 0.8408 0.8111 0.000 0.1743 0.00402 0.00061 -0.0427 0.8315 0.8369 0.250 0.2007 0.00403 0.00062 -0.0423 0.8169 0.8502 0.500 0.2257 0.00414 0.00060 -0.0416 0.7757 0.8594 0.750 0.2506 0.00432 0.00059 -0.0410 0.7212 0.8657 1.000 0.2735 0.00477 0.00063 -0.0400 0.6064 0.8715 1.250 0.2922 0.00596 0.00090 -0.0388 0.3552 0.8773 1.500 0.3141 0.00683 0.00112 -0.0382 0.1662 0.8835 1.750 0.3386 0.00733 0.00130 -0.0378 0.0664 0.8893 2.000 0.3645 0.00757 0.00144 -0.0376 0.0340 0.8951 2.250 0.3912 0.00773 0.00163 -0.0374 0.0247 0.9009 2.500 0.4173 0.00790 0.00179 -0.0372 0.0179 0.9067 2.750 0.4433 0.00816 0.00210 -0.0368 0.0139 0.9133 3.000 0.4692 0.00830 0.00228 -0.0365 0.0136 0.9193 3.250 0.4954 0.00846 0.00247 -0.0363 0.0128 0.9258 3.500 0.5209 0.00862 0.00267 -0.0360 0.0112 0.9323 3.750 0.5464 0.00887 0.00294 -0.0355 0.0089 0.9398 4.000 0.5703 0.00927 0.00345 -0.0347 0.0080 0.9479 4.250 0.5940 0.00983 0.00413 -0.0339 0.0077 0.9573 4.500 0.6185 0.01052 0.00502 -0.0333 0.0072 0.9685 4.750 0.6432 0.01241 0.00714 -0.0329 0.0057 0.9845 5.000 0.6763 0.01133 0.00594 -0.0345 0.0051 1.0000 5.250 0.7055 0.01089 0.00538 -0.0352 0.0036 1.0000 5.500 0.7325 0.01101 0.00545 -0.0353 0.0024 1.0000 5.750 0.7571 0.01167 0.00619 -0.0349 0.0017 1.0000 6.000 0.7819 0.01228 0.00689 -0.0346 0.0015 1.0000 6.250 0.8065 0.01299 0.00771 -0.0342 0.0013 1.0000 6.500 0.8310 0.01369 0.00852 -0.0338 0.0010 1.0000 6.750 0.8562 0.01424 0.00919 -0.0336 0.0008 1.0000 7.000 0.8798 0.01534 0.01045 -0.0331 0.0006 1.0000 7.250 0.9013 0.01749 0.01295 -0.0321 0.0005 1.0000 7.500 0.8472 0.05103 0.04886 -0.0181 0.0006 1.0000 7.750 0.8474 0.05761 0.05568 -0.0168 0.0007 1.0000 8.000 0.8447 0.06385 0.06210 -0.0162 0.0007 1.0000 8.250 0.8403 0.06940 0.06777 -0.0162 0.0007 1.0000 8.500 0.8319 0.07480 0.07327 -0.0169 0.0007 1.0000 8.750 0.8199 0.07881 0.07734 -0.0170 0.0007 1.0000 9.000 0.8028 0.08390 0.08246 -0.0195 0.0008 1.0000