XFOIL Version 6.96 Calculated polar for: NACA M13 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.4796 0.10991 0.10775 0.0071 1.0000 0.0151 -8.500 -0.4756 0.10656 0.10442 0.0044 1.0000 0.0151 -8.250 -0.4713 0.10309 0.10097 0.0017 1.0000 0.0152 -8.000 -0.4665 0.09950 0.09741 -0.0009 1.0000 0.0152 -7.750 -0.4605 0.09575 0.09367 -0.0041 1.0000 0.0152 -7.500 -0.4500 0.09151 0.08944 -0.0084 1.0000 0.0153 -7.250 -0.4460 0.08554 0.08349 -0.0112 1.0000 0.0155 -7.000 -0.4374 0.08225 0.08021 -0.0106 1.0000 0.0159 -6.750 -0.4253 0.07897 0.07691 -0.0126 1.0000 0.0163 -6.500 -0.4111 0.07548 0.07342 -0.0157 1.0000 0.0168 -6.250 -0.3947 0.07193 0.06985 -0.0192 1.0000 0.0177 -6.000 -0.3744 0.06811 0.06600 -0.0235 1.0000 0.0195 -5.750 -0.3411 0.06388 0.06166 -0.0309 1.0000 0.0207 -5.500 -0.3171 0.05973 0.05744 -0.0344 1.0000 0.0209 -5.250 -0.2937 0.05546 0.05308 -0.0373 1.0000 0.0209 -5.000 -0.2621 0.05041 0.04789 -0.0417 0.9876 0.0210 -4.750 -0.2354 0.04278 0.04008 -0.0469 0.9646 0.0219 -4.500 -0.2103 0.04071 0.03792 -0.0485 0.9379 0.0229 -4.250 -0.1881 0.03876 0.03581 -0.0487 0.9118 0.0248 -4.000 -0.1552 0.03647 0.03321 -0.0490 0.8901 0.0288 -3.750 -0.1285 0.03334 0.02977 -0.0491 0.8711 0.0291 -3.500 -0.1074 0.02696 0.02296 -0.0499 0.8547 0.0306 -3.250 -0.0843 0.02596 0.02185 -0.0499 0.8367 0.0322 -3.000 -0.0584 0.02444 0.02011 -0.0498 0.8202 0.0356 -2.750 -0.0306 0.02154 0.01676 -0.0496 0.8056 0.0416 -2.500 -0.0050 0.02083 0.01595 -0.0496 0.7905 0.0446 -2.250 0.0225 0.01940 0.01420 -0.0495 0.7769 0.0527 -1.500 0.1086 0.01258 0.00606 -0.0481 0.7424 0.0406 -1.250 0.1366 0.01205 0.00540 -0.0479 0.7311 0.0399 -1.000 0.1644 0.01132 0.00454 -0.0478 0.7206 0.0399 -0.750 0.1920 0.01081 0.00393 -0.0476 0.7105 0.0392 -0.500 0.2195 0.01015 0.00318 -0.0474 0.7003 0.0389 -0.250 0.2471 0.00973 0.00269 -0.0473 0.6904 0.0391 0.000 0.2748 0.00940 0.00228 -0.0472 0.6809 0.0398 0.250 0.3026 0.00916 0.00198 -0.0471 0.6711 0.0409 0.500 0.3306 0.00902 0.00180 -0.0471 0.6612 0.0420 0.750 0.3586 0.00895 0.00169 -0.0471 0.6518 0.0433 1.000 0.3866 0.00891 0.00162 -0.0471 0.6421 0.0464 1.250 0.4147 0.00887 0.00161 -0.0472 0.6317 0.0500 1.500 0.4428 0.00886 0.00162 -0.0472 0.6217 0.0633 1.750 0.4707 0.00891 0.00164 -0.0472 0.6116 0.0730 2.000 0.4987 0.00891 0.00165 -0.0472 0.6008 0.0828 2.250 0.5267 0.00891 0.00167 -0.0473 0.5894 0.0931 2.500 0.5546 0.00891 0.00172 -0.0473 0.5778 0.1102 2.750 0.5814 0.00697 0.00186 -0.0476 0.5668 1.0000 3.000 0.6088 0.00709 0.00191 -0.0475 0.5556 1.0000 3.250 0.6364 0.00719 0.00199 -0.0475 0.5443 1.0000 3.500 0.6639 0.00730 0.00208 -0.0475 0.5332 1.0000 3.750 0.6914 0.00743 0.00220 -0.0474 0.5224 1.0000 4.000 0.7188 0.00756 0.00231 -0.0474 0.5081 1.0000 4.250 0.7460 0.00771 0.00241 -0.0473 0.4897 1.0000 4.500 0.7727 0.00791 0.00252 -0.0472 0.4617 1.0000 4.750 0.7994 0.00815 0.00268 -0.0471 0.4323 1.0000 5.000 0.8258 0.00845 0.00287 -0.0470 0.3973 1.0000 5.250 0.8509 0.00901 0.00313 -0.0468 0.3253 1.0000 5.500 0.8648 0.01196 0.00455 -0.0461 0.0410 1.0000 5.750 0.8893 0.01277 0.00537 -0.0456 0.0256 1.0000 6.000 0.9147 0.01329 0.00598 -0.0452 0.0230 1.0000 6.250 0.9392 0.01393 0.00667 -0.0449 0.0200 1.0000 6.500 0.9597 0.01531 0.00819 -0.0439 0.0177 1.0000 6.750 0.9825 0.01616 0.00913 -0.0432 0.0170 1.0000 7.000 1.0046 0.01707 0.01013 -0.0424 0.0163 1.0000 7.250 1.0254 0.01821 0.01138 -0.0414 0.0155 1.0000 7.500 1.0465 0.01928 0.01252 -0.0405 0.0146 1.0000 7.750 1.0681 0.02018 0.01345 -0.0398 0.0135 1.0000 8.000 1.0881 0.02147 0.01480 -0.0388 0.0130 1.0000 8.250 1.1073 0.02315 0.01655 -0.0376 0.0125 1.0000 8.500 1.1265 0.02552 0.01905 -0.0365 0.0121 1.0000 8.750 1.1463 0.02833 0.02204 -0.0354 0.0120 1.0000 9.000 1.1665 0.03047 0.02439 -0.0342 0.0122 1.0000 18.250 0.7392 0.19971 0.19766 -0.0625 0.0123 1.0000 18.500 0.7423 0.20248 0.20044 -0.0630 0.0118 1.0000