NACA M5 AIRFOIL (m5-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M5 AIRFOIL (m5-il) Reynolds number: 100,000 Max Cl/Cd: 47.35 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m5-il-100000.txt Download as CSV file: xf-m5-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M5 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6010 0.11432 0.10963 0.0209 1.0000 0.0836
-9.000 -0.6125 0.11112 0.10651 0.0136 1.0000 0.0851
-8.750 -0.6262 0.10755 0.10298 0.0062 1.0000 0.0855
-8.500 -0.5954 0.10161 0.09702 0.0156 1.0000 0.0901
-8.250 -0.5920 0.09783 0.09325 0.0140 1.0000 0.0935
-8.000 -0.5958 0.09374 0.08920 0.0098 1.0000 0.0966
-7.750 -0.6088 0.09013 0.08549 -0.0012 1.0000 0.0996
-7.500 -0.6025 0.08422 0.07959 -0.0017 1.0000 0.1013
-7.250 -0.5878 0.08040 0.07584 0.0014 1.0000 0.1046
-7.000 -0.5792 0.07661 0.07199 -0.0005 1.0000 0.1101
-6.750 -0.5754 0.07179 0.06694 -0.0059 1.0000 0.1160
-6.500 -0.5603 0.06842 0.06367 -0.0037 1.0000 0.1221
-6.250 -0.5502 0.06420 0.05920 -0.0070 1.0000 0.1314
-6.000 -0.5344 0.06084 0.05587 -0.0062 1.0000 0.1377
-5.750 -0.5201 0.05698 0.05182 -0.0077 1.0000 0.1477
-5.500 -0.5037 0.05384 0.04843 -0.0087 1.0000 0.1609
-5.250 -0.4866 0.05082 0.04532 -0.0086 1.0000 0.1758
-5.000 -0.4689 0.04779 0.04227 -0.0083 1.0000 0.1917
-4.750 -0.4502 0.04492 0.03936 -0.0080 1.0000 0.2080
-4.500 -0.4308 0.04238 0.03676 -0.0077 1.0000 0.2278
-4.250 -0.4128 0.03983 0.03417 -0.0071 1.0000 0.2546
-4.000 -0.3602 0.03282 0.02507 -0.0099 1.0000 0.1298
-3.750 -0.3332 0.02884 0.02073 -0.0095 1.0000 0.1122
-3.500 -0.3095 0.02640 0.01804 -0.0089 1.0000 0.1082
-3.250 -0.2905 0.02480 0.01618 -0.0075 1.0000 0.1076
-3.000 -0.2772 0.02381 0.01499 -0.0054 1.0000 0.1094
-2.750 -0.2638 0.02286 0.01385 -0.0035 1.0000 0.1096
-2.500 -0.2349 0.02166 0.01238 -0.0043 0.9962 0.1099
-2.250 -0.1854 0.02026 0.01072 -0.0084 0.9866 0.1124
-2.000 -0.1364 0.01894 0.00939 -0.0128 0.9770 0.1192
-1.750 -0.0863 0.01801 0.00837 -0.0172 0.9677 0.1298
-1.500 -0.0408 0.01694 0.00745 -0.0207 0.9566 0.1433
-1.250 -0.0026 0.01604 0.00678 -0.0227 0.9425 0.1709
-1.000 0.1065 0.01316 0.00646 -0.0366 0.9490 1.0000
-0.750 0.1405 0.01324 0.00632 -0.0377 0.9275 1.0000
-0.500 0.1650 0.01338 0.00629 -0.0368 0.9050 1.0000
-0.250 0.1865 0.01353 0.00630 -0.0351 0.8858 1.0000
0.000 0.2069 0.01368 0.00632 -0.0333 0.8683 1.0000
0.250 0.2283 0.01384 0.00637 -0.0317 0.8512 1.0000
0.500 0.2498 0.01400 0.00642 -0.0301 0.8354 1.0000
0.750 0.2716 0.01416 0.00649 -0.0287 0.8203 1.0000
1.000 0.2938 0.01432 0.00657 -0.0273 0.8060 1.0000
1.250 0.3162 0.01449 0.00668 -0.0259 0.7921 1.0000
1.500 0.3388 0.01466 0.00679 -0.0246 0.7787 1.0000
1.750 0.3615 0.01484 0.00690 -0.0233 0.7659 1.0000
2.000 0.3852 0.01503 0.00706 -0.0223 0.7521 1.0000
2.250 0.4090 0.01524 0.00727 -0.0214 0.7385 1.0000
2.500 0.4329 0.01545 0.00747 -0.0204 0.7250 1.0000
2.750 0.4569 0.01568 0.00770 -0.0195 0.7117 1.0000
3.000 0.4808 0.01590 0.00793 -0.0186 0.6983 1.0000
3.250 0.5048 0.01613 0.00819 -0.0176 0.6851 1.0000
3.500 0.5288 0.01636 0.00844 -0.0166 0.6718 1.0000
3.750 0.5528 0.01658 0.00869 -0.0156 0.6583 1.0000
4.000 0.5768 0.01679 0.00893 -0.0146 0.6443 1.0000
4.250 0.6008 0.01697 0.00918 -0.0135 0.6289 1.0000
4.500 0.6247 0.01713 0.00939 -0.0124 0.6125 1.0000
4.750 0.6486 0.01725 0.00956 -0.0112 0.5952 1.0000
5.000 0.6717 0.01713 0.00944 -0.0094 0.5749 1.0000
5.250 0.6946 0.01691 0.00923 -0.0077 0.5477 1.0000
5.500 0.7178 0.01669 0.00902 -0.0061 0.5186 1.0000
5.750 0.7412 0.01656 0.00891 -0.0047 0.4874 1.0000
6.000 0.7647 0.01649 0.00895 -0.0035 0.4473 1.0000
6.250 0.7864 0.01661 0.00893 -0.0021 0.3711 1.0000
6.500 0.7946 0.01991 0.01049 -0.0006 0.1294 1.0000
6.750 0.8116 0.02183 0.01218 0.0007 0.0968 1.0000
7.000 0.8297 0.02347 0.01375 0.0021 0.0842 1.0000
7.250 0.8477 0.02530 0.01541 0.0035 0.0755 1.0000
7.500 0.8696 0.02669 0.01692 0.0047 0.0688 1.0000
7.750 0.8918 0.02880 0.01891 0.0057 0.0647 1.0000
8.000 0.9157 0.03125 0.02149 0.0066 0.0621 1.0000
8.250 0.9381 0.03314 0.02369 0.0077 0.0586 1.0000
8.500 0.9598 0.03557 0.02640 0.0087 0.0567 1.0000
8.750 0.9796 0.03852 0.02977 0.0098 0.0562 1.0000
9.000 0.9956 0.04200 0.03376 0.0112 0.0568 1.0000
9.250 1.0075 0.04588 0.03815 0.0127 0.0577 1.0000
9.500 1.0151 0.04980 0.04254 0.0142 0.0580 1.0000
9.750 1.0179 0.05394 0.04713 0.0156 0.0584 1.0000
10.000 1.0164 0.05849 0.05206 0.0170 0.0596 1.0000
10.250 1.0146 0.06353 0.05734 0.0179 0.0616 1.0000
10.500 0.9852 0.06903 0.06346 0.0193 0.0679 1.0000
10.750 0.9516 0.07418 0.06885 0.0192 0.0699 1.0000
11.000 0.9243 0.08045 0.07525 0.0164 0.0721 1.0000
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Polar data table (+)
Polar graphs
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