XFOIL Version 6.94 Calculated polar for: NACA M7 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2157 0.01457 0.00728 -0.0088 0.5583 0.0268 0.500 0.2725 0.01240 0.00472 -0.0082 0.5495 0.0267 1.500 0.3843 0.01089 0.00311 -0.0077 0.5235 0.0434 2.000 0.4401 0.01054 0.00279 -0.0076 0.5190 0.0592 2.500 0.4957 0.01018 0.00254 -0.0075 0.5035 0.1270 3.000 0.6102 0.00862 0.00300 -0.0206 0.4933 1.0000 3.500 0.6638 0.00875 0.00300 -0.0202 0.4806 1.0000 4.000 0.7170 0.00885 0.00306 -0.0197 0.4657 1.0000 4.500 0.7702 0.00901 0.00319 -0.0193 0.4528 1.0000 5.000 0.8228 0.00916 0.00327 -0.0189 0.4317 1.0000 5.500 0.8757 0.00937 0.00346 -0.0185 0.4072 1.0000 6.000 0.9282 0.00965 0.00374 -0.0182 0.3891 1.0000 6.500 0.9797 0.01023 0.00412 -0.0180 0.3292 1.0000 7.000 1.0198 0.01423 0.00663 -0.0190 0.0632 1.0000 7.500 1.0665 0.01544 0.00779 -0.0183 0.0531 1.0000 8.000 1.1123 0.01652 0.00896 -0.0176 0.0504 1.0000 8.500 1.1548 0.01788 0.01042 -0.0166 0.0474 1.0000 9.000 1.1931 0.01957 0.01224 -0.0152 0.0469 1.0000 9.500 1.2262 0.02145 0.01420 -0.0137 0.0451 1.0000 10.000 1.2497 0.02402 0.01689 -0.0115 0.0434 1.0000 10.500 1.2702 0.02645 0.01946 -0.0099 0.0435 1.0000 11.000 1.2850 0.02978 0.02294 -0.0083 0.0414 1.0000 11.500 1.3027 0.03293 0.02612 -0.0071 0.0398 1.0000 12.000 1.3246 0.03559 0.02882 -0.0043 0.0395 1.0000 12.500 1.3426 0.03893 0.03243 -0.0042 0.0378 1.0000 13.000 1.3609 0.04219 0.03579 -0.0036 0.0358 1.0000 13.500 1.3842 0.04486 0.03859 -0.0019 0.0357 1.0000 14.000 1.4085 0.04763 0.04153 0.0003 0.0347 1.0000 15.500 1.4218 0.06232 0.05694 -0.0018 0.0301 1.0000 16.000 1.4204 0.06859 0.06349 -0.0040 0.0286 1.0000 16.500 1.4105 0.07716 0.07229 -0.0089 0.0261 1.0000 17.000 1.4106 0.08365 0.07905 -0.0106 0.0254 1.0000 17.500 1.4080 0.09163 0.08716 -0.0149 0.0238 1.0000 18.500 1.4020 0.10988 0.10586 -0.0252 0.0092 1.0000 19.000 1.3923 0.11869 0.11492 -0.0285 0.0095 1.0000 19.500 1.3628 0.13138 0.12780 -0.0340 0.0077 1.0000 20.000 1.3353 0.14449 0.14115 -0.0403 0.0071 1.0000 21.000 1.2824 0.17242 0.16957 -0.0554 0.0070 1.0000