XFOIL Version 6.96 Calculated polar for: NACA M25 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 ------ -------- --------- --------- -------- -------- -------- -10.500 -0.2940 0.12857 0.12568 0.0204 0.6413 0.0092 -10.250 -0.2892 0.12553 0.12260 0.0195 0.6376 0.0097 -7.750 -0.3396 0.10804 0.10495 0.0200 0.6132 0.0106 -7.500 -0.3331 0.10398 0.10084 0.0196 0.6095 0.0107 -7.250 -0.3250 0.10089 0.09773 0.0180 0.6057 0.0109 -7.000 -0.3136 0.09780 0.09464 0.0156 0.6015 0.0110 -6.750 -0.3005 0.09474 0.09156 0.0130 0.5976 0.0113 -6.500 -0.2861 0.09168 0.08844 0.0102 0.5942 0.0116 -6.250 -0.2701 0.08855 0.08528 0.0073 0.5906 0.0119 -6.000 -0.2527 0.08536 0.08208 0.0043 0.5867 0.0124 -5.750 -0.2338 0.08213 0.07881 0.0011 0.5830 0.0130 -5.500 -0.2107 0.07902 0.07562 -0.0026 0.5795 0.0136 -5.250 -0.1814 0.07623 0.07277 -0.0076 0.5760 0.0138 -5.000 -0.1503 0.07343 0.06988 -0.0128 0.5724 0.0139 -4.500 -0.1169 0.06537 0.06173 -0.0156 0.5656 0.0143 -4.250 -0.0952 0.06265 0.05895 -0.0175 0.5621 0.0150 -4.000 -0.0696 0.05993 0.05618 -0.0200 0.5584 0.0159 -3.750 -0.0418 0.05711 0.05328 -0.0226 0.5550 0.0168 -3.500 -0.0061 0.05490 0.05093 -0.0258 0.5516 0.0177 -3.250 0.0322 0.05290 0.04875 -0.0293 0.5483 0.0179 -3.000 0.0670 0.05039 0.04611 -0.0318 0.5447 0.0180 -2.750 0.0858 0.04591 0.04159 -0.0326 0.5416 0.0183 -2.500 0.1057 0.04327 0.03890 -0.0333 0.5383 0.0189 -2.250 0.1323 0.04130 0.03681 -0.0344 0.5349 0.0199 -2.000 0.1621 0.03920 0.03464 -0.0357 0.5316 0.0212 -1.750 0.2044 0.03843 0.03365 -0.0372 0.5280 0.0230 -1.500 0.2368 0.03662 0.03168 -0.0381 0.5248 0.0232 -1.250 0.2674 0.03477 0.02964 -0.0387 0.5217 0.0233 -1.000 0.2881 0.03093 0.02573 -0.0393 0.5186 0.0241 -0.750 0.3138 0.02927 0.02402 -0.0399 0.5151 0.0249 -0.500 0.3420 0.02785 0.02249 -0.0403 0.5118 0.0262 0.500 0.4649 0.02219 0.01617 -0.0411 0.4989 0.0316 0.750 0.4913 0.02109 0.01500 -0.0414 0.4956 0.0327 1.000 0.5190 0.02024 0.01401 -0.0417 0.4925 0.0342 1.250 0.5480 0.01941 0.01310 -0.0418 0.4894 0.0368 1.500 0.5797 0.01923 0.01277 -0.0416 0.4858 0.0403 1.750 0.6080 0.01731 0.01065 -0.0416 0.4825 0.0420 2.000 0.6351 0.01667 0.00995 -0.0419 0.4792 0.0438 2.250 0.6632 0.01617 0.00937 -0.0421 0.4760 0.0474 2.500 0.6925 0.01541 0.00847 -0.0421 0.4727 0.0550 2.750 0.7202 0.01503 0.00810 -0.0424 0.4692 0.0602 3.000 0.7486 0.01454 0.00749 -0.0425 0.4660 0.0690 3.750 0.8345 0.01268 0.00530 -0.0423 0.4557 0.0552 4.000 0.8617 0.01294 0.00569 -0.0430 0.4522 0.0818 4.250 0.8888 0.01187 0.00435 -0.0420 0.4493 0.0423 4.500 0.9161 0.01172 0.00424 -0.0420 0.4460 0.0418 4.750 0.9431 0.01155 0.00410 -0.0420 0.4424 0.0430 5.000 0.9705 0.01149 0.00405 -0.0422 0.4381 0.0453 5.250 0.9980 0.01147 0.00406 -0.0424 0.4314 0.0480 5.500 1.0258 0.01144 0.00405 -0.0426 0.4238 0.0519 5.750 1.0535 0.01141 0.00405 -0.0429 0.4126 0.0592 6.000 1.0807 0.01138 0.00419 -0.0432 0.4015 0.1809 6.500 1.1752 0.01149 0.00509 -0.0540 0.2972 1.0000 6.750 1.1826 0.01644 0.00862 -0.0555 0.0277 1.0000 7.000 1.2048 0.01702 0.00928 -0.0551 0.0236 1.0000 7.250 1.2255 0.01776 0.01011 -0.0547 0.0209 1.0000 7.500 1.2433 0.01882 0.01133 -0.0540 0.0182 1.0000 7.750 1.2579 0.02009 0.01274 -0.0532 0.0173 1.0000 8.000 1.2721 0.02123 0.01397 -0.0524 0.0168 1.0000 8.250 1.2816 0.02280 0.01565 -0.0516 0.0163 1.0000 8.500 1.2818 0.02510 0.01806 -0.0509 0.0159 1.0000 8.750 1.2764 0.02760 0.02067 -0.0492 0.0157 1.0000 9.000 1.2737 0.03036 0.02352 -0.0484 0.0154 1.0000 9.250 1.2717 0.03321 0.02646 -0.0479 0.0149 1.0000 9.500 1.2693 0.03614 0.02949 -0.0474 0.0143 1.0000 9.750 1.2664 0.03914 0.03255 -0.0469 0.0138 1.0000 10.000 1.2622 0.04228 0.03577 -0.0463 0.0136 1.0000 10.250 1.2578 0.04541 0.03897 -0.0457 0.0134 1.0000 10.500 1.2541 0.04842 0.04205 -0.0450 0.0132 1.0000 10.750 1.2514 0.05124 0.04493 -0.0442 0.0130 1.0000 11.000 1.2507 0.05384 0.04757 -0.0433 0.0128 1.0000