NACA M12 AIRFOIL (m12-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M12 AIRFOIL (m12-il) Reynolds number: 50,000 Max Cl/Cd: 31.6 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m12-il-50000.txt Download as CSV file: xf-m12-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4846 0.12858 0.12107 -0.0028 1.0000 0.2510
-10.250 -0.4688 0.12405 0.11654 -0.0024 1.0000 0.2597
-10.000 -0.4970 0.12374 0.11636 -0.0048 1.0000 0.2689
-9.750 -0.4631 0.11792 0.11048 -0.0031 1.0000 0.2814
-9.500 -0.4582 0.11433 0.10692 -0.0033 1.0000 0.2904
-9.250 -0.4773 0.11287 0.10558 -0.0045 1.0000 0.3023
-9.000 -0.4565 0.10867 0.10135 -0.0035 1.0000 0.3160
-8.750 -0.4420 0.10475 0.09743 -0.0030 1.0000 0.3281
-8.500 -0.4362 0.10161 0.09434 -0.0025 1.0000 0.3436
-8.250 -0.4310 0.09868 0.09145 -0.0017 1.0000 0.3617
-8.000 -0.4311 0.09613 0.08896 -0.0008 1.0000 0.3813
-7.750 -0.4484 0.09460 0.08756 0.0003 1.0000 0.4007
-7.500 -0.4205 0.09011 0.08304 0.0014 1.0000 0.4198
-7.250 -0.4044 0.08673 0.07967 0.0024 1.0000 0.4392
-7.000 -0.3945 0.08384 0.07681 0.0037 1.0000 0.4606
-6.750 -0.3937 0.08137 0.07442 0.0057 1.0000 0.4859
-6.250 -0.5622 0.05636 0.04862 -0.0237 1.0000 0.1865
-6.000 -0.5599 0.05172 0.04361 -0.0222 1.0000 0.1737
-5.750 -0.5503 0.04850 0.04012 -0.0203 1.0000 0.1672
-5.500 -0.5423 0.04517 0.03644 -0.0184 1.0000 0.1617
-5.250 -0.5343 0.04235 0.03267 -0.0159 1.0000 0.1554
-5.000 -0.5202 0.03998 0.02998 -0.0141 1.0000 0.1553
-4.750 -0.5046 0.03764 0.02753 -0.0126 1.0000 0.1573
-4.500 -0.4876 0.03576 0.02552 -0.0112 1.0000 0.1599
-4.250 -0.4691 0.03401 0.02349 -0.0099 1.0000 0.1618
-4.000 -0.4491 0.03239 0.02159 -0.0087 1.0000 0.1642
-3.750 -0.4279 0.03098 0.01987 -0.0076 1.0000 0.1683
-3.500 -0.4065 0.02966 0.01840 -0.0067 1.0000 0.1755
-3.250 -0.3841 0.02861 0.01721 -0.0059 1.0000 0.1857
-3.000 -0.3594 0.02744 0.01609 -0.0053 1.0000 0.1980
-2.750 -0.1047 0.02243 0.01432 -0.0342 1.0000 1.0000
-2.500 -0.1128 0.02241 0.01413 -0.0299 1.0000 1.0000
-2.250 -0.1191 0.02246 0.01402 -0.0257 1.0000 1.0000
-2.000 -0.1233 0.02256 0.01396 -0.0217 1.0000 1.0000
-1.750 -0.1252 0.02272 0.01395 -0.0181 1.0000 1.0000
-1.500 -0.1249 0.02292 0.01399 -0.0148 1.0000 1.0000
-1.250 -0.1225 0.02318 0.01408 -0.0119 1.0000 1.0000
-1.000 -0.1182 0.02349 0.01421 -0.0094 1.0000 1.0000
-0.750 -0.1124 0.02385 0.01441 -0.0071 1.0000 1.0000
-0.500 -0.1049 0.02426 0.01468 -0.0052 1.0000 1.0000
-0.250 -0.0955 0.02475 0.01502 -0.0036 0.9999 1.0000
0.000 -0.0410 0.02568 0.01573 -0.0104 0.9858 1.0000
0.250 0.0099 0.02658 0.01645 -0.0163 0.9717 1.0000
0.500 0.0579 0.02739 0.01713 -0.0215 0.9571 1.0000
0.750 0.1019 0.02815 0.01778 -0.0259 0.9423 1.0000
1.000 0.1435 0.02886 0.01841 -0.0296 0.9272 1.0000
1.250 0.1823 0.02956 0.01904 -0.0327 0.9121 1.0000
1.500 0.2191 0.03024 0.01968 -0.0353 0.8969 1.0000
1.750 0.2545 0.03092 0.02034 -0.0375 0.8818 1.0000
2.000 0.2894 0.03158 0.02099 -0.0395 0.8668 1.0000
2.250 0.3245 0.03223 0.02164 -0.0414 0.8519 1.0000
2.500 0.3606 0.03283 0.02226 -0.0433 0.8371 1.0000
2.750 0.3986 0.03336 0.02283 -0.0454 0.8222 1.0000
3.000 0.4393 0.03377 0.02330 -0.0477 0.8073 1.0000
3.250 0.4795 0.03404 0.02364 -0.0494 0.7926 1.0000
3.500 0.5150 0.03430 0.02398 -0.0502 0.7774 1.0000
3.750 0.5429 0.03470 0.02445 -0.0499 0.7615 1.0000
4.000 0.5697 0.03510 0.02492 -0.0493 0.7455 1.0000
4.250 0.5960 0.03549 0.02540 -0.0485 0.7295 1.0000
4.500 0.6218 0.03586 0.02586 -0.0475 0.7135 1.0000
4.750 0.6475 0.03620 0.02629 -0.0464 0.6974 1.0000
5.000 0.6733 0.03651 0.02670 -0.0452 0.6812 1.0000
5.250 0.7002 0.03668 0.02698 -0.0440 0.6647 1.0000
5.500 0.7277 0.03676 0.02718 -0.0426 0.6482 1.0000
5.750 0.7575 0.03661 0.02716 -0.0413 0.6314 1.0000
6.000 0.7790 0.03696 0.02762 -0.0393 0.6128 1.0000
6.250 0.7965 0.03752 0.02829 -0.0370 0.5926 1.0000
6.500 0.8262 0.03703 0.02794 -0.0351 0.5724 1.0000
6.750 0.8701 0.03515 0.02612 -0.0336 0.5511 1.0000
7.000 0.8894 0.03512 0.02619 -0.0309 0.5257 1.0000
7.250 0.9227 0.03377 0.02485 -0.0286 0.4984 1.0000
7.500 0.9511 0.03272 0.02374 -0.0261 0.4678 1.0000
7.750 0.9721 0.03220 0.02318 -0.0232 0.4339 1.0000
8.000 0.9905 0.03181 0.02270 -0.0200 0.3958 1.0000
8.250 1.0046 0.03179 0.02248 -0.0165 0.3526 1.0000
8.500 1.0147 0.03239 0.02275 -0.0128 0.3060 1.0000
8.750 1.0248 0.03365 0.02359 -0.0096 0.2633 1.0000
9.000 1.0362 0.03545 0.02516 -0.0070 0.2303 1.0000
9.250 1.0552 0.03736 0.02678 -0.0055 0.2056 1.0000
9.500 1.0709 0.03949 0.02894 -0.0038 0.1881 1.0000
9.750 1.0907 0.04173 0.03116 -0.0028 0.1744 1.0000
10.000 1.1041 0.04407 0.03375 -0.0010 0.1643 1.0000
10.250 1.1163 0.04676 0.03667 0.0007 0.1563 1.0000
10.500 1.1323 0.04936 0.03943 0.0020 0.1493 1.0000
10.750 1.1398 0.05257 0.04290 0.0038 0.1445 1.0000
11.000 1.1346 0.05591 0.04669 0.0067 0.1411 1.0000
11.250 1.1324 0.05923 0.05029 0.0091 0.1378 1.0000
11.500 1.1517 0.06247 0.05350 0.0095 0.1334 1.0000
11.750 1.1390 0.06654 0.05786 0.0120 0.1323 1.0000
12.000 1.1159 0.07056 0.06219 0.0151 0.1321 1.0000
12.250 1.0900 0.07488 0.06673 0.0175 0.1323 1.0000
12.500 1.0636 0.07994 0.07197 0.0184 0.1326 1.0000
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Polar data table (+)
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