NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 50,000 Max Cl/Cd: 34.01 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m4-il-50000-n5.txt Download as CSV file: xf-m4-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M4 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 -0.6048 0.12399 0.11750 0.0265 1.0000 0.0808
-9.000 -0.6092 0.12156 0.11515 0.0219 1.0000 0.0818
-8.750 -0.6134 0.11878 0.11245 0.0168 1.0000 0.0822
-8.500 -0.5910 0.11156 0.10519 0.0229 1.0000 0.0876
-8.250 -0.5877 0.10788 0.10155 0.0209 1.0000 0.0908
-8.000 -0.5896 0.10463 0.09838 0.0168 1.0000 0.0943
-7.750 -0.5933 0.10204 0.09580 0.0088 1.0000 0.0961
-7.500 -0.5857 0.09685 0.09065 0.0077 1.0000 0.0975
-7.250 -0.5743 0.09200 0.08583 0.0095 1.0000 0.1006
-7.000 -0.5657 0.08795 0.08177 0.0076 1.0000 0.1042
-6.500 -0.5475 0.07972 0.07349 0.0022 1.0000 0.1160
-6.250 -0.5364 0.07619 0.06980 -0.0028 1.0000 0.1262
-6.000 -0.5232 0.07186 0.06547 -0.0016 1.0000 0.1339
-5.750 -0.4904 0.06312 0.05609 -0.0090 1.0000 0.0666
-5.500 -0.4736 0.05860 0.05146 -0.0097 1.0000 0.0634
-5.250 -0.4537 0.05409 0.04668 -0.0109 1.0000 0.0612
-5.000 -0.4322 0.05014 0.04240 -0.0119 1.0000 0.0635
-4.750 -0.4091 0.04622 0.03804 -0.0126 1.0000 0.0653
-4.500 -0.3856 0.04241 0.03382 -0.0129 1.0000 0.0649
-4.250 -0.3607 0.03889 0.02980 -0.0130 1.0000 0.0656
-4.000 -0.3364 0.03608 0.02658 -0.0130 1.0000 0.0709
-3.750 -0.3108 0.03347 0.02357 -0.0129 1.0000 0.0727
-3.500 -0.2838 0.03095 0.02057 -0.0127 1.0000 0.0733
-3.250 -0.2564 0.02878 0.01799 -0.0125 1.0000 0.0745
-3.000 -0.2287 0.02693 0.01577 -0.0123 1.0000 0.0765
-2.750 -0.2011 0.02561 0.01403 -0.0121 1.0000 0.0834
-2.500 -0.1742 0.02424 0.01253 -0.0121 1.0000 0.0869
-2.250 -0.1459 0.02300 0.01110 -0.0120 1.0000 0.0882
-2.000 -0.1163 0.02198 0.00989 -0.0122 1.0000 0.0899
-1.750 -0.0894 0.02117 0.00894 -0.0121 1.0000 0.0921
-1.500 -0.0663 0.02057 0.00817 -0.0114 1.0000 0.0947
-1.250 -0.0467 0.02010 0.00765 -0.0104 1.0000 0.0980
-1.000 -0.0078 0.01948 0.00697 -0.0131 0.9828 0.1067
-0.750 0.0336 0.01882 0.00635 -0.0161 0.9641 0.1284
-0.500 0.1066 0.01573 0.00584 -0.0247 0.9598 1.0000
-0.250 0.1477 0.01586 0.00562 -0.0276 0.9378 1.0000
0.000 0.1833 0.01601 0.00552 -0.0292 0.9151 1.0000
0.250 0.2151 0.01620 0.00552 -0.0299 0.8940 1.0000
0.500 0.2427 0.01642 0.00559 -0.0297 0.8733 1.0000
0.750 0.2685 0.01664 0.00568 -0.0290 0.8546 1.0000
1.000 0.2927 0.01688 0.00583 -0.0282 0.8357 1.0000
1.250 0.3166 0.01713 0.00600 -0.0271 0.8182 1.0000
1.500 0.3402 0.01738 0.00621 -0.0260 0.8016 1.0000
1.750 0.3637 0.01763 0.00642 -0.0249 0.7858 1.0000
2.000 0.3872 0.01790 0.00666 -0.0238 0.7704 1.0000
2.250 0.4110 0.01817 0.00697 -0.0228 0.7550 1.0000
2.500 0.4349 0.01846 0.00729 -0.0219 0.7396 1.0000
2.750 0.4589 0.01876 0.00764 -0.0210 0.7245 1.0000
3.000 0.4831 0.01906 0.00803 -0.0201 0.7094 1.0000
3.250 0.5073 0.01938 0.00848 -0.0192 0.6943 1.0000
3.500 0.5317 0.01971 0.00893 -0.0183 0.6791 1.0000
3.750 0.5560 0.02005 0.00941 -0.0175 0.6636 1.0000
4.000 0.5804 0.02038 0.00991 -0.0166 0.6479 1.0000
4.250 0.6045 0.02070 0.01046 -0.0155 0.6314 1.0000
4.500 0.6283 0.02095 0.01089 -0.0142 0.6137 1.0000
4.750 0.6522 0.02123 0.01141 -0.0131 0.5917 1.0000
5.000 0.6744 0.02124 0.01159 -0.0111 0.5615 1.0000
5.250 0.6915 0.02074 0.01097 -0.0075 0.4890 1.0000
5.500 0.7111 0.02091 0.01099 -0.0053 0.4059 1.0000
5.750 0.7295 0.02194 0.01150 -0.0037 0.2576 1.0000
6.000 0.7423 0.02520 0.01348 -0.0033 0.1118 1.0000
6.250 0.7601 0.02737 0.01553 -0.0024 0.0868 1.0000
6.500 0.7780 0.02925 0.01750 -0.0015 0.0696 1.0000
6.750 0.7958 0.03123 0.01957 -0.0001 0.0621 1.0000
7.000 0.8153 0.03303 0.02153 0.0012 0.0539 1.0000
7.250 0.8361 0.03526 0.02387 0.0026 0.0490 1.0000
7.500 0.8593 0.03759 0.02652 0.0040 0.0455 1.0000
7.750 0.8796 0.04000 0.02912 0.0048 0.0413 1.0000
8.000 0.8985 0.04324 0.03260 0.0057 0.0386 1.0000
8.250 0.9161 0.04664 0.03656 0.0067 0.0376 1.0000
8.500 0.9296 0.05042 0.04098 0.0077 0.0369 1.0000
8.750 0.9385 0.05439 0.04548 0.0085 0.0361 1.0000
9.000 0.9425 0.05847 0.05003 0.0092 0.0352 1.0000
9.250 0.9418 0.06264 0.05460 0.0097 0.0344 1.0000
9.500 0.9363 0.06691 0.05921 0.0098 0.0340 1.0000
9.750 0.9257 0.07134 0.06391 0.0095 0.0338 1.0000
10.000 0.9090 0.07584 0.06859 0.0088 0.0341 1.0000
10.250 0.8909 0.08123 0.07412 0.0060 0.0346 1.0000
10.500 0.8734 0.08757 0.08055 0.0019 0.0353 1.0000
10.750 0.8579 0.09458 0.08761 -0.0029 0.0361 1.0000
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Polar data table (+)
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