NACA M6 AIRFOIL (m6-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M6 AIRFOIL (m6-il) Reynolds number: 50,000 Max Cl/Cd: 32 at α=8.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m6-il-50000-n5.txt Download as CSV file: xf-m6-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5707 0.08588 0.07860 -0.0277 1.0000 0.0726
-9.000 -0.5767 0.08175 0.07446 -0.0282 1.0000 0.0723
-8.750 -0.5842 0.07736 0.07001 -0.0288 1.0000 0.0720
-8.500 -0.5914 0.07293 0.06547 -0.0291 1.0000 0.0720
-8.250 -0.5970 0.06859 0.06096 -0.0288 1.0000 0.0721
-8.000 -0.6001 0.06444 0.05658 -0.0282 1.0000 0.0722
-7.750 -0.5998 0.06053 0.05240 -0.0271 1.0000 0.0723
-7.500 -0.5961 0.05692 0.04852 -0.0259 1.0000 0.0722
-7.250 -0.5900 0.05354 0.04487 -0.0244 1.0000 0.0721
-7.000 -0.5822 0.05041 0.04146 -0.0228 1.0000 0.0720
-6.750 -0.5735 0.04752 0.03827 -0.0209 1.0000 0.0720
-6.500 -0.5646 0.04491 0.03535 -0.0188 1.0000 0.0722
-6.250 -0.5559 0.04260 0.03273 -0.0164 1.0000 0.0724
-6.000 -0.5477 0.04058 0.03034 -0.0137 1.0000 0.0730
-5.750 -0.5385 0.03896 0.02864 -0.0114 1.0000 0.0742
-5.500 -0.5292 0.03772 0.02732 -0.0090 1.0000 0.0756
-5.250 -0.5053 0.03618 0.02559 -0.0093 0.9942 0.0776
-5.000 -0.4673 0.03416 0.02321 -0.0119 0.9834 0.0795
-4.750 -0.4282 0.03226 0.02095 -0.0144 0.9727 0.0813
-4.500 -0.3879 0.03060 0.01895 -0.0169 0.9624 0.0836
-4.250 -0.3455 0.02915 0.01721 -0.0197 0.9528 0.0866
-4.000 -0.3033 0.02802 0.01613 -0.0227 0.9437 0.0922
-3.750 -0.2629 0.02707 0.01498 -0.0250 0.9324 0.0993
-3.500 -0.2222 0.02610 0.01403 -0.0274 0.9221 0.1069
-3.250 -0.1828 0.02519 0.01308 -0.0295 0.9120 0.1192
-3.000 -0.1507 0.02441 0.01233 -0.0304 0.8992 0.1435
-2.750 -0.1220 0.02350 0.01170 -0.0308 0.8869 0.1869
-2.500 -0.0988 0.02239 0.01121 -0.0303 0.8749 0.2851
-2.250 -0.0045 0.02143 0.01252 -0.0378 0.8744 0.9287
-2.000 0.1126 0.02142 0.01195 -0.0543 0.8701 1.0000
-1.750 0.1362 0.02141 0.01169 -0.0537 0.8546 1.0000
-1.500 0.1591 0.02141 0.01148 -0.0530 0.8401 1.0000
-1.250 0.1818 0.02143 0.01130 -0.0521 0.8266 1.0000
-1.000 0.2043 0.02146 0.01114 -0.0512 0.8141 1.0000
-0.750 0.2267 0.02152 0.01105 -0.0503 0.8012 1.0000
-0.500 0.2492 0.02161 0.01100 -0.0495 0.7887 1.0000
-0.250 0.2718 0.02169 0.01095 -0.0486 0.7774 1.0000
0.000 0.2943 0.02177 0.01090 -0.0476 0.7668 1.0000
0.250 0.3167 0.02192 0.01097 -0.0468 0.7550 1.0000
0.500 0.3392 0.02206 0.01101 -0.0459 0.7446 1.0000
0.750 0.3616 0.02219 0.01105 -0.0449 0.7344 1.0000
1.000 0.3839 0.02241 0.01122 -0.0441 0.7233 1.0000
1.250 0.4065 0.02255 0.01128 -0.0430 0.7141 1.0000
1.500 0.4288 0.02278 0.01149 -0.0422 0.7033 1.0000
1.750 0.4514 0.02303 0.01170 -0.0414 0.6932 1.0000
2.000 0.4742 0.02319 0.01182 -0.0403 0.6841 1.0000
2.250 0.4961 0.02352 0.01217 -0.0395 0.6731 1.0000
2.500 0.5187 0.02372 0.01234 -0.0384 0.6642 1.0000
2.750 0.5406 0.02402 0.01266 -0.0374 0.6537 1.0000
3.000 0.5623 0.02435 0.01301 -0.0364 0.6434 1.0000
3.250 0.5850 0.02453 0.01317 -0.0352 0.6346 1.0000
3.500 0.6058 0.02494 0.01365 -0.0343 0.6232 1.0000
3.750 0.6278 0.02521 0.01395 -0.0331 0.6136 1.0000
4.000 0.6495 0.02548 0.01426 -0.0319 0.6034 1.0000
4.250 0.6700 0.02589 0.01474 -0.0307 0.5920 1.0000
4.500 0.6922 0.02611 0.01499 -0.0295 0.5822 1.0000
4.750 0.7133 0.02641 0.01536 -0.0282 0.5710 1.0000
5.000 0.7332 0.02683 0.01587 -0.0269 0.5591 1.0000
5.250 0.7546 0.02709 0.01621 -0.0256 0.5482 1.0000
5.500 0.7767 0.02727 0.01645 -0.0242 0.5375 1.0000
5.750 0.7957 0.02774 0.01704 -0.0228 0.5247 1.0000
6.000 0.8156 0.02812 0.01754 -0.0214 0.5126 1.0000
6.250 0.8371 0.02835 0.01785 -0.0200 0.5013 1.0000
6.500 0.8584 0.02859 0.01819 -0.0186 0.4895 1.0000
6.750 0.8767 0.02908 0.01882 -0.0171 0.4761 1.0000
7.000 0.8954 0.02951 0.01940 -0.0155 0.4626 1.0000
7.250 0.9143 0.02975 0.01974 -0.0137 0.4472 1.0000
7.500 0.9327 0.02984 0.01990 -0.0117 0.4293 1.0000
7.750 0.9512 0.02976 0.01981 -0.0096 0.4093 1.0000
8.000 0.9634 0.03022 0.02041 -0.0072 0.3869 1.0000
8.250 0.9776 0.03055 0.02078 -0.0049 0.3654 1.0000
8.500 0.9890 0.03107 0.02133 -0.0025 0.3418 1.0000
8.750 0.9984 0.03162 0.02184 0.0002 0.3159 1.0000
9.000 1.0041 0.03251 0.02271 0.0030 0.2882 1.0000
9.250 1.0075 0.03359 0.02375 0.0059 0.2607 1.0000
9.500 1.0081 0.03490 0.02496 0.0089 0.2343 1.0000
9.750 1.0043 0.03647 0.02640 0.0121 0.2107 1.0000
10.000 0.9993 0.03835 0.02819 0.0149 0.1890 1.0000
10.250 0.9946 0.04053 0.03027 0.0170 0.1697 1.0000
10.500 0.9908 0.04292 0.03257 0.0186 0.1533 1.0000
10.750 0.9884 0.04540 0.03500 0.0199 0.1395 1.0000
11.000 0.9877 0.04790 0.03746 0.0210 0.1283 1.0000
11.250 0.9875 0.05043 0.03995 0.0218 0.1193 1.0000
11.500 0.9887 0.05291 0.04243 0.0226 0.1112 1.0000
11.750 0.9919 0.05531 0.04487 0.0234 0.1044 1.0000
12.000 0.9960 0.05762 0.04719 0.0241 0.0987 1.0000
12.250 1.0014 0.05992 0.04953 0.0249 0.0935 1.0000
12.500 1.0064 0.06236 0.05210 0.0255 0.0885 1.0000
12.750 1.0151 0.06428 0.05395 0.0264 0.0843 1.0000
13.000 1.0204 0.06701 0.05694 0.0269 0.0807 1.0000
13.250 1.0226 0.07004 0.06019 0.0270 0.0773 1.0000
13.500 1.0272 0.07267 0.06291 0.0272 0.0741 1.0000
13.750 1.0384 0.07476 0.06498 0.0281 0.0712 1.0000
14.000 1.0301 0.07931 0.06986 0.0272 0.0697 1.0000
14.250 1.0188 0.08436 0.07521 0.0257 0.0684 1.0000
14.500 1.0042 0.09004 0.08116 0.0237 0.0673 1.0000
14.750 0.9858 0.09658 0.08794 0.0210 0.0665 1.0000
15.000 0.9614 0.10467 0.09626 0.0170 0.0662 1.0000
15.250 0.9240 0.11633 0.10817 0.0105 0.0668 1.0000
15.500 0.8738 0.13300 0.12497 0.0009 0.0681 1.0000
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