XFOIL Version 6.96 Calculated polar for: NACA 0006 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.7042 0.09180 0.08516 0.0035 1.0000 0.0398 -8.500 -0.7055 0.08682 0.08020 0.0010 1.0000 0.0395 -8.250 -0.7061 0.08170 0.07504 -0.0017 1.0000 0.0392 -8.000 -0.7060 0.07655 0.06982 -0.0041 1.0000 0.0389 -7.750 -0.7048 0.07139 0.06454 -0.0063 1.0000 0.0386 -7.500 -0.7016 0.06624 0.05920 -0.0081 1.0000 0.0383 -7.250 -0.6960 0.06114 0.05384 -0.0095 1.0000 0.0380 -7.000 -0.6878 0.05610 0.04845 -0.0104 1.0000 0.0378 -6.750 -0.6766 0.05127 0.04318 -0.0108 1.0000 0.0378 -6.500 -0.6623 0.04669 0.03808 -0.0108 1.0000 0.0380 -6.250 -0.6451 0.04252 0.03320 -0.0104 1.0000 0.0393 -6.000 -0.6270 0.03891 0.02922 -0.0100 1.0000 0.0422 -5.750 -0.6057 0.03598 0.02593 -0.0095 1.0000 0.0452 -5.500 -0.5820 0.03284 0.02223 -0.0086 1.0000 0.0475 -5.250 -0.5569 0.03021 0.01896 -0.0078 1.0000 0.0528 -5.000 -0.5329 0.02801 0.01661 -0.0071 1.0000 0.0599 -4.750 -0.5072 0.02591 0.01423 -0.0061 1.0000 0.0680 -4.500 -0.4825 0.02429 0.01246 -0.0053 1.0000 0.0817 -4.250 -0.4574 0.02271 0.01069 -0.0043 1.0000 0.0938 -4.000 -0.4327 0.02137 0.00934 -0.0038 1.0000 0.1202 -3.750 -0.4087 0.01978 0.00791 -0.0033 1.0000 0.1583 -3.500 -0.3881 0.01792 0.00685 -0.0025 1.0000 0.2999 -3.250 -0.3731 0.01634 0.00646 0.0003 1.0000 0.5394 -3.000 -0.3543 0.01551 0.00658 0.0053 1.0000 0.7856 -2.750 -0.2956 0.01553 0.00629 0.0021 1.0000 0.9329 -2.500 -0.2247 0.01532 0.00551 -0.0062 1.0000 0.9906 -2.250 -0.1933 0.01503 0.00489 -0.0078 1.0000 1.0000 -2.000 -0.1723 0.01481 0.00445 -0.0071 1.0000 1.0000 -1.750 -0.1511 0.01464 0.00409 -0.0063 1.0000 1.0000 -1.500 -0.1297 0.01450 0.00379 -0.0055 1.0000 1.0000 -1.250 -0.1081 0.01439 0.00352 -0.0047 1.0000 1.0000 -1.000 -0.0866 0.01431 0.00332 -0.0038 1.0000 1.0000 -0.750 -0.0650 0.01424 0.00318 -0.0029 1.0000 1.0000 -0.500 -0.0434 0.01420 0.00307 -0.0019 1.0000 1.0000 -0.250 -0.0217 0.01417 0.00300 -0.0010 1.0000 1.0000 0.000 0.0000 0.01416 0.00298 0.0000 1.0000 1.0000 0.250 0.0217 0.01417 0.00300 0.0010 1.0000 1.0000 0.500 0.0434 0.01420 0.00307 0.0019 1.0000 1.0000 0.750 0.0650 0.01424 0.00317 0.0028 1.0000 1.0000 1.000 0.0866 0.01430 0.00332 0.0038 1.0000 1.0000 1.250 0.1082 0.01439 0.00352 0.0047 1.0000 1.0000 1.500 0.1297 0.01450 0.00379 0.0055 1.0000 1.0000 1.750 0.1512 0.01464 0.00409 0.0063 1.0000 1.0000 2.000 0.1724 0.01481 0.00445 0.0071 1.0000 1.0000 2.250 0.1934 0.01503 0.00488 0.0077 1.0000 1.0000 2.500 0.2247 0.01532 0.00551 0.0062 0.9907 1.0000 2.750 0.2956 0.01553 0.00629 -0.0021 0.9330 1.0000 3.000 0.3543 0.01551 0.00658 -0.0053 0.7856 1.0000 3.250 0.3731 0.01634 0.00646 -0.0003 0.5397 1.0000 3.500 0.3881 0.01792 0.00685 0.0025 0.3002 1.0000 3.750 0.4087 0.01978 0.00791 0.0033 0.1583 1.0000 4.000 0.4327 0.02137 0.00934 0.0038 0.1202 1.0000 4.250 0.4574 0.02271 0.01069 0.0043 0.0938 1.0000 4.500 0.4825 0.02429 0.01246 0.0053 0.0817 1.0000 4.750 0.5072 0.02591 0.01423 0.0061 0.0680 1.0000 5.000 0.5329 0.02801 0.01662 0.0071 0.0599 1.0000 5.250 0.5570 0.03022 0.01896 0.0077 0.0528 1.0000 5.500 0.5821 0.03284 0.02224 0.0086 0.0475 1.0000 5.750 0.6058 0.03598 0.02593 0.0095 0.0452 1.0000 6.000 0.6270 0.03891 0.02922 0.0100 0.0421 1.0000 6.250 0.6451 0.04252 0.03321 0.0104 0.0393 1.0000 6.500 0.6624 0.04670 0.03809 0.0108 0.0380 1.0000 6.750 0.6767 0.05128 0.04319 0.0108 0.0378 1.0000 7.000 0.6880 0.05611 0.04846 0.0103 0.0378 1.0000 7.250 0.6962 0.06115 0.05385 0.0094 0.0380 1.0000 7.500 0.7018 0.06626 0.05922 0.0081 0.0383 1.0000 7.750 0.7050 0.07141 0.06456 0.0063 0.0386 1.0000 8.000 0.7063 0.07658 0.06985 0.0041 0.0389 1.0000 8.250 0.7064 0.08173 0.07507 0.0016 0.0392 1.0000 8.500 0.7059 0.08685 0.08023 -0.0010 0.0395 1.0000 8.750 0.7046 0.09184 0.08520 -0.0036 0.0398 1.0000