XFOIL Version 6.96 Calculated polar for: NACA 0008 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.000 -0.5788 0.11151 0.10819 0.0018 1.0000 0.0557 -10.750 -0.5877 0.10591 0.10262 -0.0012 1.0000 0.0558 -8.250 -0.7847 0.05406 0.04942 -0.0151 1.0000 0.0433 -8.000 -0.7592 0.02794 0.02274 -0.0161 1.0000 0.0331 -7.750 -0.7840 0.03747 0.03154 -0.0120 1.0000 0.0326 -7.250 -0.7586 0.02796 0.02099 -0.0092 1.0000 0.0340 -7.000 -0.7376 0.02636 0.01935 -0.0086 1.0000 0.0371 -6.750 -0.7152 0.02460 0.01732 -0.0077 1.0000 0.0401 -6.500 -0.6916 0.02301 0.01533 -0.0066 1.0000 0.0428 -6.250 -0.6697 0.02064 0.01282 -0.0058 1.0000 0.0475 -6.000 -0.6455 0.01954 0.01160 -0.0051 1.0000 0.0523 -5.750 -0.6214 0.01829 0.01018 -0.0042 1.0000 0.0573 -5.500 -0.5979 0.01717 0.00908 -0.0035 1.0000 0.0625 -5.250 -0.5732 0.01652 0.00832 -0.0028 1.0000 0.0677 -5.000 -0.5502 0.01541 0.00719 -0.0019 1.0000 0.0721 -4.750 -0.5265 0.01470 0.00649 -0.0011 1.0000 0.0781 -4.500 -0.5033 0.01392 0.00569 -0.0002 1.0000 0.0853 -4.250 -0.4800 0.01321 0.00500 0.0007 1.0000 0.0971 -4.000 -0.4586 0.01210 0.00421 0.0017 1.0000 0.1433 -3.750 -0.4386 0.01088 0.00371 0.0026 1.0000 0.2836 -3.500 -0.4170 0.01017 0.00344 0.0036 1.0000 0.3863 -3.250 -0.3950 0.00962 0.00325 0.0046 1.0000 0.4752 -3.000 -0.3731 0.00914 0.00310 0.0058 1.0000 0.5617 -2.750 -0.3511 0.00873 0.00302 0.0073 1.0000 0.6412 -2.500 -0.3293 0.00841 0.00300 0.0089 1.0000 0.7171 -2.250 -0.3078 0.00819 0.00302 0.0109 1.0000 0.7860 -2.000 -0.2861 0.00810 0.00310 0.0130 1.0000 0.8479 -1.750 -0.2608 0.00815 0.00322 0.0145 1.0000 0.8999 -1.500 -0.2250 0.00828 0.00333 0.0137 1.0000 0.9393 -1.250 -0.1781 0.00842 0.00339 0.0101 1.0000 0.9653 -1.000 -0.1228 0.00850 0.00339 0.0045 1.0000 0.9806 -0.750 -0.0661 0.00851 0.00333 -0.0016 1.0000 0.9922 -0.500 -0.0189 0.00845 0.00323 -0.0059 1.0000 1.0000 -0.250 -0.0087 0.00835 0.00313 -0.0031 1.0000 1.0000 0.000 0.0000 0.00832 0.00310 0.0000 1.0000 1.0000 0.250 0.0087 0.00835 0.00313 0.0031 1.0000 1.0000 0.500 0.0189 0.00845 0.00323 0.0059 1.0000 1.0000 0.750 0.0660 0.00851 0.00333 0.0016 0.9922 1.0000 1.000 0.1228 0.00850 0.00339 -0.0045 0.9806 1.0000 1.250 0.1780 0.00842 0.00339 -0.0101 0.9653 1.0000 1.500 0.2250 0.00828 0.00333 -0.0137 0.9393 1.0000 1.750 0.2608 0.00815 0.00323 -0.0145 0.9000 1.0000 2.000 0.2861 0.00810 0.00310 -0.0130 0.8480 1.0000 2.250 0.3078 0.00819 0.00302 -0.0109 0.7860 1.0000 2.500 0.3293 0.00841 0.00300 -0.0089 0.7171 1.0000 2.750 0.3511 0.00873 0.00302 -0.0073 0.6412 1.0000 3.000 0.3731 0.00914 0.00310 -0.0058 0.5616 1.0000 3.250 0.3950 0.00962 0.00325 -0.0046 0.4751 1.0000 3.500 0.4170 0.01017 0.00344 -0.0036 0.3863 1.0000 3.750 0.4387 0.01088 0.00371 -0.0026 0.2836 1.0000 4.000 0.4586 0.01209 0.00421 -0.0017 0.1434 1.0000 4.250 0.4800 0.01321 0.00500 -0.0007 0.0971 1.0000 4.500 0.5033 0.01392 0.00569 0.0002 0.0853 1.0000 4.750 0.5266 0.01470 0.00649 0.0011 0.0781 1.0000 5.000 0.5502 0.01541 0.00719 0.0019 0.0721 1.0000 5.250 0.5732 0.01652 0.00832 0.0028 0.0677 1.0000 5.500 0.5979 0.01717 0.00908 0.0035 0.0625 1.0000 5.750 0.6214 0.01829 0.01018 0.0042 0.0573 1.0000 6.000 0.6455 0.01954 0.01160 0.0051 0.0523 1.0000 6.250 0.6697 0.02064 0.01283 0.0058 0.0475 1.0000 6.500 0.6916 0.02300 0.01532 0.0066 0.0428 1.0000 6.750 0.7152 0.02460 0.01731 0.0077 0.0401 1.0000 7.000 0.7376 0.02635 0.01934 0.0086 0.0371 1.0000 7.250 0.7585 0.02796 0.02099 0.0092 0.0340 1.0000 7.750 0.7840 0.03746 0.03153 0.0120 0.0326 1.0000 8.250 0.7379 0.03815 0.03411 0.0194 0.0397 1.0000