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 = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.7405 0.08222 0.08077 0.0149 1.0000 0.0070 -8.500 -0.7508 0.07216 0.07073 0.0032 1.0000 0.0069 -8.250 -0.7533 0.06306 0.06151 -0.0039 1.0000 0.0068 -8.000 -0.7521 0.05597 0.05427 -0.0074 1.0000 0.0069 -7.750 -0.7462 0.05056 0.04869 -0.0090 1.0000 0.0070 -7.500 -0.7370 0.04574 0.04365 -0.0098 1.0000 0.0072 -7.250 -0.7382 0.03375 0.03109 -0.0092 1.0000 0.0056 -7.000 -0.7345 0.02389 0.02039 -0.0071 1.0000 0.0052 -6.750 -0.7169 0.01969 0.01565 -0.0058 1.0000 0.0053 -6.500 -0.6954 0.01711 0.01268 -0.0048 1.0000 0.0054 -6.250 -0.6709 0.01593 0.01132 -0.0043 1.0000 0.0058 -6.000 -0.6461 0.01480 0.01001 -0.0037 1.0000 0.0061 -5.750 -0.6209 0.01377 0.00881 -0.0031 1.0000 0.0063 -5.500 -0.6001 0.01111 0.00580 -0.0018 1.0000 0.0071 -5.250 -0.5746 0.01047 0.00509 -0.0014 1.0000 0.0079 -5.000 -0.5488 0.00994 0.00451 -0.0010 1.0000 0.0088 -4.750 -0.5225 0.00959 0.00410 -0.0007 1.0000 0.0099 -4.500 -0.4973 0.00878 0.00313 0.0000 1.0000 0.0128 -4.250 -0.4711 0.00839 0.00275 0.0004 1.0000 0.0182 -4.000 -0.4448 0.00812 0.00250 0.0007 1.0000 0.0281 -3.750 -0.4184 0.00796 0.00233 0.0010 1.0000 0.0330 -3.500 -0.3923 0.00774 0.00211 0.0013 1.0000 0.0377 -3.250 -0.3662 0.00761 0.00196 0.0016 1.0000 0.0400 -3.000 -0.3403 0.00737 0.00170 0.0021 1.0000 0.0446 -2.750 -0.3145 0.00718 0.00152 0.0025 1.0000 0.0498 -2.500 -0.2890 0.00689 0.00133 0.0029 1.0000 0.0753 -2.250 -0.2645 0.00622 0.00113 0.0033 1.0000 0.1947 -2.000 -0.2396 0.00570 0.00100 0.0036 1.0000 0.3040 -1.750 -0.2142 0.00530 0.00090 0.0038 1.0000 0.3937 -1.500 -0.1884 0.00495 0.00085 0.0041 1.0000 0.4785 -1.250 -0.1608 0.00464 0.00081 0.0039 0.9995 0.5581 -1.000 -0.1251 0.00426 0.00077 0.0020 0.9958 0.6541 -0.750 -0.0896 0.00395 0.00074 0.0002 0.9904 0.7336 -0.500 -0.0529 0.00365 0.00073 -0.0016 0.9776 0.8123 -0.250 -0.0237 0.00344 0.00073 -0.0015 0.9539 0.8808 0.000 0.0000 0.00338 0.00074 0.0000 0.9238 0.9241 0.250 0.0238 0.00344 0.00073 0.0015 0.8810 0.9537 0.500 0.0528 0.00365 0.00073 0.0016 0.8123 0.9774 0.750 0.0896 0.00396 0.00074 -0.0002 0.7330 0.9905 1.000 0.1251 0.00426 0.00077 -0.0020 0.6539 0.9958 1.250 0.1609 0.00464 0.00081 -0.0039 0.5581 0.9995 1.500 0.1883 0.00495 0.00085 -0.0040 0.4790 1.0000 1.750 0.2140 0.00530 0.00090 -0.0038 0.3938 1.0000 2.000 0.2394 0.00570 0.00100 -0.0035 0.3043 1.0000 2.250 0.2643 0.00622 0.00113 -0.0032 0.1959 1.0000 2.500 0.2888 0.00689 0.00133 -0.0029 0.0756 1.0000 2.750 0.3143 0.00718 0.00152 -0.0024 0.0499 1.0000 3.000 0.3401 0.00737 0.00170 -0.0020 0.0447 1.0000 3.250 0.3660 0.00761 0.00196 -0.0016 0.0400 1.0000 3.500 0.3921 0.00774 0.00211 -0.0012 0.0377 1.0000 3.750 0.4182 0.00796 0.00233 -0.0009 0.0331 1.0000 4.000 0.4446 0.00812 0.00250 -0.0006 0.0281 1.0000 4.250 0.4709 0.00839 0.00274 -0.0003 0.0181 1.0000 4.500 0.4972 0.00878 0.00313 0.0001 0.0127 1.0000 4.750 0.5225 0.00958 0.00409 0.0007 0.0099 1.0000 5.000 0.5487 0.00996 0.00452 0.0010 0.0088 1.0000 5.250 0.5746 0.01046 0.00509 0.0014 0.0079 1.0000 5.500 0.6001 0.01111 0.00581 0.0018 0.0071 1.0000 5.750 0.6211 0.01371 0.00874 0.0031 0.0063 1.0000 6.000 0.6462 0.01479 0.01000 0.0037 0.0061 1.0000 6.250 0.6710 0.01594 0.01133 0.0043 0.0058 1.0000 6.500 0.6954 0.01713 0.01270 0.0048 0.0054 1.0000 6.750 0.7170 0.01971 0.01568 0.0058 0.0053 1.0000 7.000 0.7348 0.02384 0.02034 0.0071 0.0052 1.0000 7.250 0.7383 0.03381 0.03116 0.0091 0.0056 1.0000 10.500 0.7164 0.12067 0.11910 -0.0360 0.0063 1.0000 10.750 0.7174 0.12546 0.12388 -0.0379 0.0063 1.0000