XFOIL Version 6.94 Calculated polar for: NACA M18 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.2697 0.01909 0.01313 -0.0368 0.5855 0.1862 1.500 0.3988 0.01350 0.00576 -0.0335 0.5642 0.0475 2.000 0.4506 0.01284 0.00492 -0.0327 0.5514 0.0447 2.500 0.5033 0.01262 0.00445 -0.0322 0.5342 0.0432 3.000 0.5552 0.01335 0.00505 -0.0322 0.5000 0.0436 3.500 0.6101 0.01321 0.00483 -0.0324 0.4969 0.0483 4.000 0.7444 0.01139 0.00540 -0.0499 0.4884 0.9927 4.500 0.8219 0.01148 0.00542 -0.0553 0.4752 1.0000 5.000 0.8719 0.01161 0.00543 -0.0550 0.4605 1.0000 5.500 0.9215 0.01188 0.00559 -0.0547 0.4468 1.0000 6.000 0.9701 0.01232 0.00591 -0.0542 0.4252 1.0000 6.500 1.0180 0.01325 0.00690 -0.0540 0.4003 1.0000 7.000 1.0672 0.01308 0.00675 -0.0534 0.3979 1.0000 7.500 1.1145 0.01285 0.00648 -0.0525 0.3821 1.0000 8.000 1.1566 0.01331 0.00663 -0.0510 0.3395 1.0000 8.500 1.1953 0.01407 0.00724 -0.0490 0.2966 1.0000 9.000 1.2258 0.01535 0.00804 -0.0459 0.2372 1.0000 9.500 1.2232 0.01819 0.01003 -0.0380 0.1374 1.0000 10.000 1.2084 0.02094 0.01243 -0.0282 0.0753 1.0000 10.500 1.1973 0.02413 0.01542 -0.0211 0.0458 1.0000 11.000 1.1675 0.03058 0.02231 -0.0154 0.0181 1.0000 11.500 1.1544 0.03673 0.02876 -0.0133 0.0168 1.0000 12.000 1.1385 0.04341 0.03571 -0.0120 0.0166 1.0000 12.500 1.1331 0.04906 0.04156 -0.0104 0.0167 1.0000 13.000 1.1328 0.05426 0.04697 -0.0086 0.0172 1.0000 13.500 1.1354 0.05911 0.05216 -0.0054 0.0188 1.0000 14.000 1.1333 0.06500 0.05865 0.0001 0.0224 1.0000