NACA M3 AIRFOIL (m3-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M3 AIRFOIL (m3-il) Reynolds number: 100,000 Max Cl/Cd: 37.94 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m3-il-100000.txt Download as CSV file: xf-m3-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M3 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.5320 0.11150 0.10633 -0.0127 1.0000 0.1820
-10.750 -0.8318 0.07298 0.06731 -0.0326 1.0000 0.0934
-10.500 -0.8480 0.06838 0.06261 -0.0315 1.0000 0.0920
-10.250 -0.8690 0.06367 0.05772 -0.0296 1.0000 0.0904
-10.000 -0.8915 0.05838 0.05209 -0.0272 1.0000 0.0881
-9.750 -0.9239 0.05188 0.04479 -0.0228 1.0000 0.0844
-9.500 -0.9296 0.04814 0.04039 -0.0194 1.0000 0.0832
-9.250 -0.9224 0.04504 0.03688 -0.0171 1.0000 0.0829
-9.000 -0.9114 0.04183 0.03335 -0.0153 1.0000 0.0832
-8.750 -0.8959 0.03915 0.03054 -0.0141 1.0000 0.0849
-8.500 -0.8779 0.03717 0.02846 -0.0129 1.0000 0.0869
-8.250 -0.8611 0.03508 0.02609 -0.0113 1.0000 0.0884
-8.000 -0.8430 0.03303 0.02375 -0.0097 1.0000 0.0897
-7.750 -0.8235 0.03120 0.02164 -0.0083 1.0000 0.0914
-7.500 -0.8039 0.02977 0.01991 -0.0067 1.0000 0.0941
-7.250 -0.7816 0.02790 0.01798 -0.0058 1.0000 0.0972
-7.000 -0.7589 0.02650 0.01660 -0.0049 1.0000 0.1006
-6.750 -0.7366 0.02526 0.01528 -0.0038 1.0000 0.1047
-6.500 -0.7144 0.02402 0.01395 -0.0026 1.0000 0.1098
-6.250 -0.6941 0.02289 0.01297 -0.0013 1.0000 0.1174
-6.000 -0.6747 0.02175 0.01190 0.0002 1.0000 0.1269
-5.750 -0.6573 0.02061 0.01091 0.0022 1.0000 0.1416
-5.500 -0.6425 0.01936 0.00993 0.0044 1.0000 0.1708
-5.250 -0.6303 0.01797 0.00903 0.0070 1.0000 0.2311
-5.000 -0.6179 0.01681 0.00842 0.0095 1.0000 0.3136
-4.750 -0.6033 0.01603 0.00806 0.0118 1.0000 0.3921
-4.500 -0.5871 0.01547 0.00784 0.0140 1.0000 0.4621
-4.250 -0.5701 0.01506 0.00772 0.0162 1.0000 0.5290
-4.000 -0.5528 0.01480 0.00775 0.0187 1.0000 0.5944
-3.750 -0.5347 0.01468 0.00785 0.0213 1.0000 0.6508
-3.500 -0.5156 0.01466 0.00794 0.0236 1.0000 0.6973
-3.250 -0.4955 0.01472 0.00809 0.0260 1.0000 0.7354
-3.000 -0.4758 0.01486 0.00828 0.0284 1.0000 0.7722
-2.750 -0.4563 0.01508 0.00854 0.0311 1.0000 0.8067
-2.500 -0.4359 0.01540 0.00886 0.0336 1.0000 0.8382
-2.250 -0.4124 0.01577 0.00921 0.0356 1.0000 0.8666
-2.000 -0.3852 0.01616 0.00955 0.0365 1.0000 0.8939
-1.750 -0.3429 0.01670 0.01000 0.0346 1.0000 0.9163
-1.500 -0.2910 0.01721 0.01040 0.0304 1.0000 0.9362
-1.250 -0.2263 0.01762 0.01070 0.0233 1.0000 0.9506
-1.000 -0.1623 0.01785 0.01085 0.0159 1.0000 0.9638
-0.750 -0.1009 0.01795 0.01088 0.0087 1.0000 0.9769
-0.500 -0.0391 0.01793 0.01083 0.0012 1.0000 0.9895
-0.250 0.0161 0.01784 0.01072 -0.0054 1.0000 1.0000
0.000 0.0000 0.01781 0.01070 0.0000 1.0000 1.0000
0.250 -0.0161 0.01783 0.01072 0.0054 1.0000 1.0000
0.500 0.0391 0.01793 0.01083 -0.0012 0.9895 1.0000
0.750 0.1008 0.01795 0.01088 -0.0087 0.9769 1.0000
1.000 0.1623 0.01785 0.01085 -0.0159 0.9638 1.0000
1.250 0.2263 0.01762 0.01070 -0.0233 0.9506 1.0000
1.500 0.2910 0.01720 0.01040 -0.0304 0.9363 1.0000
1.750 0.3428 0.01670 0.01000 -0.0346 0.9163 1.0000
2.000 0.3851 0.01616 0.00954 -0.0365 0.8939 1.0000
2.250 0.4124 0.01577 0.00921 -0.0355 0.8666 1.0000
2.500 0.4358 0.01539 0.00885 -0.0336 0.8383 1.0000
2.750 0.4562 0.01508 0.00854 -0.0310 0.8067 1.0000
3.000 0.4757 0.01485 0.00828 -0.0284 0.7722 1.0000
3.250 0.4954 0.01471 0.00808 -0.0260 0.7355 1.0000
3.500 0.5155 0.01466 0.00794 -0.0236 0.6973 1.0000
3.750 0.5346 0.01468 0.00785 -0.0212 0.6508 1.0000
4.000 0.5527 0.01479 0.00775 -0.0187 0.5946 1.0000
4.250 0.5700 0.01506 0.00771 -0.0162 0.5292 1.0000
4.500 0.5870 0.01547 0.00783 -0.0140 0.4622 1.0000
4.750 0.6032 0.01602 0.00806 -0.0117 0.3922 1.0000
5.000 0.6178 0.01680 0.00842 -0.0094 0.3138 1.0000
5.250 0.6303 0.01796 0.00903 -0.0070 0.2313 1.0000
5.500 0.6424 0.01936 0.00992 -0.0044 0.1710 1.0000
5.750 0.6572 0.02061 0.01090 -0.0021 0.1417 1.0000
6.000 0.6747 0.02174 0.01190 -0.0002 0.1270 1.0000
6.250 0.6941 0.02289 0.01297 0.0013 0.1174 1.0000
6.500 0.7144 0.02402 0.01395 0.0026 0.1098 1.0000
6.750 0.7365 0.02526 0.01527 0.0038 0.1047 1.0000
7.000 0.7589 0.02649 0.01659 0.0049 0.1006 1.0000
7.250 0.7815 0.02790 0.01797 0.0058 0.0972 1.0000
7.500 0.8038 0.02977 0.01990 0.0067 0.0941 1.0000
7.750 0.8235 0.03120 0.02164 0.0083 0.0914 1.0000
8.000 0.8430 0.03302 0.02375 0.0097 0.0897 1.0000
8.250 0.8611 0.03508 0.02609 0.0113 0.0884 1.0000
8.500 0.8780 0.03716 0.02845 0.0129 0.0869 1.0000
8.750 0.8960 0.03914 0.03053 0.0140 0.0849 1.0000
9.000 0.9114 0.04184 0.03335 0.0152 0.0832 1.0000
9.250 0.9224 0.04506 0.03690 0.0171 0.0829 1.0000
9.500 0.9297 0.04815 0.04040 0.0193 0.0832 1.0000
9.750 0.9232 0.05194 0.04486 0.0229 0.0845 1.0000
10.000 0.8916 0.05841 0.05212 0.0271 0.0881 1.0000
10.250 0.8692 0.06370 0.05775 0.0295 0.0904 1.0000
10.500 0.8483 0.06841 0.06265 0.0315 0.0920 1.0000
10.750 0.8322 0.07304 0.06736 0.0325 0.0934 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M3 AIRFOIL (m3-il)