NACA M6 AIRFOIL (m6-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M6 AIRFOIL (m6-il) Reynolds number: 100,000 Max Cl/Cd: 48.85 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m6-il-100000-n5.txt Download as CSV file: xf-m6-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M6 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5594 0.09139 0.08627 -0.0157 1.0000 0.0490
-9.500 -0.5707 0.08433 0.07924 -0.0215 1.0000 0.0486
-9.250 -0.5896 0.07695 0.07183 -0.0269 1.0000 0.0481
-9.000 -0.6095 0.07109 0.06589 -0.0283 1.0000 0.0478
-8.750 -0.6233 0.06549 0.06013 -0.0288 1.0000 0.0478
-8.500 -0.6329 0.06018 0.05461 -0.0285 1.0000 0.0478
-8.250 -0.6393 0.05499 0.04912 -0.0274 1.0000 0.0476
-8.000 -0.6419 0.05010 0.04388 -0.0258 1.0000 0.0473
-7.750 -0.6395 0.04592 0.03933 -0.0239 1.0000 0.0471
-7.500 -0.6333 0.04238 0.03541 -0.0220 1.0000 0.0471
-7.250 -0.6242 0.03940 0.03208 -0.0200 1.0000 0.0473
-7.000 -0.6137 0.03700 0.02938 -0.0179 1.0000 0.0477
-6.750 -0.6038 0.03520 0.02731 -0.0154 1.0000 0.0485
-6.500 -0.5752 0.03283 0.02452 -0.0166 0.9867 0.0497
-6.250 -0.5432 0.03053 0.02179 -0.0180 0.9730 0.0505
-6.000 -0.5104 0.02854 0.01943 -0.0194 0.9597 0.0510
-5.750 -0.4775 0.02688 0.01742 -0.0206 0.9461 0.0515
-5.500 -0.4452 0.02548 0.01574 -0.0215 0.9323 0.0522
-5.250 -0.4143 0.02425 0.01437 -0.0221 0.9183 0.0529
-5.000 -0.3855 0.02328 0.01335 -0.0224 0.9041 0.0541
-4.750 -0.3584 0.02257 0.01259 -0.0222 0.8900 0.0560
-4.500 -0.3321 0.02189 0.01181 -0.0217 0.8762 0.0581
-4.250 -0.3065 0.02120 0.01099 -0.0210 0.8632 0.0598
-4.000 -0.2816 0.02055 0.01023 -0.0202 0.8497 0.0613
-3.750 -0.2575 0.01989 0.00956 -0.0193 0.8367 0.0631
-3.500 -0.2334 0.01935 0.00899 -0.0184 0.8246 0.0655
-3.250 -0.2092 0.01888 0.00845 -0.0175 0.8135 0.0689
-3.000 -0.1847 0.01844 0.00794 -0.0166 0.8018 0.0738
-2.750 -0.1603 0.01798 0.00749 -0.0158 0.7907 0.0830
-2.500 -0.1361 0.01751 0.00708 -0.0149 0.7810 0.1002
-2.250 -0.1114 0.01703 0.00677 -0.0143 0.7702 0.1356
-2.000 -0.0865 0.01660 0.00649 -0.0136 0.7605 0.1837
-1.750 -0.0625 0.01605 0.00626 -0.0129 0.7511 0.2657
-1.500 -0.0100 0.01446 0.00707 -0.0158 0.7426 0.8614
-1.250 0.0308 0.01511 0.00755 -0.0167 0.7338 0.9278
-1.000 0.0932 0.01563 0.00787 -0.0227 0.7238 0.9617
-0.750 0.1486 0.01580 0.00782 -0.0280 0.7149 0.9814
-0.500 0.2152 0.01574 0.00759 -0.0360 0.7042 1.0000
-0.250 0.2398 0.01570 0.00741 -0.0356 0.6951 1.0000
0.250 0.2894 0.01565 0.00716 -0.0348 0.6769 1.0000
0.500 0.3140 0.01564 0.00704 -0.0343 0.6683 1.0000
0.750 0.3392 0.01565 0.00699 -0.0339 0.6591 1.0000
1.000 0.3639 0.01566 0.00690 -0.0334 0.6509 1.0000
1.250 0.3891 0.01570 0.00691 -0.0331 0.6415 1.0000
1.500 0.4138 0.01573 0.00684 -0.0325 0.6338 1.0000
1.750 0.4390 0.01579 0.00690 -0.0322 0.6241 1.0000
2.000 0.4637 0.01584 0.00687 -0.0316 0.6164 1.0000
2.250 0.4886 0.01591 0.00696 -0.0313 0.6066 1.0000
2.500 0.5131 0.01599 0.00699 -0.0307 0.5985 1.0000
2.750 0.5376 0.01608 0.00709 -0.0301 0.5889 1.0000
3.000 0.5620 0.01618 0.00717 -0.0296 0.5800 1.0000
3.250 0.5862 0.01627 0.00725 -0.0289 0.5708 1.0000
3.500 0.6104 0.01639 0.00740 -0.0283 0.5607 1.0000
3.750 0.6344 0.01649 0.00746 -0.0276 0.5517 1.0000
4.000 0.6583 0.01663 0.00764 -0.0269 0.5405 1.0000
4.250 0.6820 0.01676 0.00781 -0.0262 0.5301 1.0000
4.500 0.7057 0.01688 0.00791 -0.0254 0.5204 1.0000
4.750 0.7291 0.01705 0.00816 -0.0247 0.5089 1.0000
5.000 0.7525 0.01723 0.00838 -0.0239 0.4982 1.0000
5.500 0.7987 0.01761 0.00884 -0.0222 0.4767 1.0000
5.750 0.8214 0.01782 0.00911 -0.0213 0.4649 1.0000
6.000 0.8438 0.01802 0.00936 -0.0203 0.4516 1.0000
6.250 0.8656 0.01823 0.00960 -0.0193 0.4366 1.0000
6.500 0.8870 0.01847 0.00987 -0.0181 0.4199 1.0000
6.750 0.9079 0.01873 0.01018 -0.0169 0.4021 1.0000
7.000 0.9278 0.01902 0.01046 -0.0156 0.3810 1.0000
7.250 0.9467 0.01938 0.01080 -0.0142 0.3558 1.0000
7.500 0.9645 0.01982 0.01117 -0.0126 0.3289 1.0000
7.750 0.9813 0.02035 0.01165 -0.0110 0.3009 1.0000
8.000 0.9976 0.02095 0.01220 -0.0093 0.2735 1.0000
8.250 1.0122 0.02166 0.01283 -0.0075 0.2455 1.0000
8.500 1.0250 0.02248 0.01356 -0.0055 0.2168 1.0000
8.750 1.0357 0.02344 0.01442 -0.0032 0.1875 1.0000
9.000 1.0441 0.02454 0.01539 -0.0007 0.1576 1.0000
9.250 1.0503 0.02576 0.01647 0.0019 0.1297 1.0000
9.750 1.0570 0.02844 0.01894 0.0077 0.0972 1.0000
10.000 1.0577 0.02975 0.02026 0.0109 0.0889 1.0000
10.250 1.0547 0.03133 0.02179 0.0140 0.0831 1.0000
10.500 1.0569 0.03283 0.02341 0.0163 0.0774 1.0000
10.750 1.0566 0.03469 0.02528 0.0182 0.0731 1.0000
11.000 1.0587 0.03655 0.02720 0.0197 0.0693 1.0000
11.250 1.0626 0.03837 0.02913 0.0209 0.0654 1.0000
11.500 1.0648 0.04042 0.03122 0.0218 0.0623 1.0000
11.750 1.0673 0.04254 0.03337 0.0228 0.0597 1.0000
12.000 1.0729 0.04446 0.03543 0.0235 0.0569 1.0000
12.250 1.0773 0.04652 0.03759 0.0241 0.0543 1.0000
12.500 1.0804 0.04874 0.03984 0.0246 0.0522 1.0000
12.750 1.0847 0.05090 0.04204 0.0252 0.0503 1.0000
13.000 1.0907 0.05298 0.04428 0.0258 0.0483 1.0000
13.250 1.0953 0.05521 0.04662 0.0261 0.0464 1.0000
13.500 1.0989 0.05756 0.04903 0.0263 0.0447 1.0000
13.750 1.1025 0.05989 0.05137 0.0265 0.0433 1.0000
14.000 1.1065 0.06236 0.05397 0.0268 0.0419 1.0000
14.250 1.1094 0.06504 0.05686 0.0270 0.0406 1.0000
14.500 1.1107 0.06791 0.05989 0.0269 0.0393 1.0000
14.750 1.1110 0.07090 0.06302 0.0266 0.0382 1.0000
15.000 1.1114 0.07389 0.06609 0.0261 0.0372 1.0000
15.250 1.1129 0.07673 0.06895 0.0257 0.0362 1.0000
15.500 1.1115 0.08016 0.07250 0.0252 0.0353 1.0000
15.750 1.1028 0.08485 0.07745 0.0236 0.0347 1.0000
16.000 1.0930 0.08983 0.08269 0.0219 0.0341 1.0000
16.250 1.0815 0.09531 0.08839 0.0197 0.0337 1.0000
16.500 1.0676 0.10139 0.09472 0.0169 0.0333 1.0000
16.750 1.0505 0.10832 0.10186 0.0134 0.0331 1.0000
17.000 1.0291 0.11648 0.11025 0.0089 0.0330 1.0000
17.250 0.9999 0.12696 0.12097 0.0027 0.0331 1.0000
17.500 0.9483 0.14414 0.13842 -0.0079 0.0337 1.0000
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