NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M2 AIRFOIL (m2-il) Reynolds number: 50,000 Max Cl/Cd: 21.82 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m2-il-50000.txt Download as CSV file: xf-m2-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M2 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.6202 0.09770 0.09068 0.0371 1.0000 0.4281
-7.750 -0.6183 0.09515 0.08816 0.0393 1.0000 0.4592
-7.500 -0.6110 0.09215 0.08518 0.0414 1.0000 0.4890
-7.250 -0.5821 0.08707 0.08004 0.0417 1.0000 0.5099
-7.000 -0.5753 0.08351 0.07649 0.0422 1.0000 0.5250
-6.750 -0.7002 0.05621 0.04829 -0.0100 1.0000 0.1807
-6.500 -0.6867 0.05000 0.04118 -0.0111 1.0000 0.1557
-6.250 -0.6682 0.04624 0.03690 -0.0106 1.0000 0.1526
-6.000 -0.6484 0.04257 0.03282 -0.0100 1.0000 0.1517
-5.750 -0.6264 0.03910 0.02890 -0.0094 1.0000 0.1494
-5.500 -0.6030 0.03609 0.02538 -0.0086 1.0000 0.1488
-5.250 -0.5789 0.03386 0.02246 -0.0077 1.0000 0.1536
-5.000 -0.5540 0.03120 0.01975 -0.0072 1.0000 0.1598
-4.750 -0.5272 0.02908 0.01728 -0.0064 1.0000 0.1650
-4.500 -0.4998 0.02698 0.01500 -0.0057 1.0000 0.1732
-4.250 -0.4731 0.02510 0.01313 -0.0050 1.0000 0.1913
-4.000 -0.4445 0.02305 0.01115 -0.0043 1.0000 0.2168
-3.750 -0.4188 0.02079 0.00930 -0.0035 1.0000 0.2668
-3.500 -0.4036 0.01850 0.00829 -0.0013 1.0000 0.3936
-3.250 -0.3865 0.01785 0.00975 0.0094 1.0000 0.8523
-3.000 -0.2063 0.01968 0.01010 -0.0121 1.0000 0.9941
-2.750 -0.1749 0.01896 0.00911 -0.0146 1.0000 1.0000
-2.500 -0.1579 0.01837 0.00837 -0.0142 1.0000 1.0000
-2.250 -0.1409 0.01786 0.00773 -0.0136 1.0000 1.0000
-2.000 -0.1239 0.01743 0.00719 -0.0129 1.0000 1.0000
-1.750 -0.1071 0.01707 0.00673 -0.0119 1.0000 1.0000
-1.500 -0.0904 0.01677 0.00634 -0.0107 1.0000 1.0000
-1.250 -0.0741 0.01653 0.00603 -0.0094 1.0000 1.0000
-1.000 -0.0583 0.01634 0.00579 -0.0078 1.0000 1.0000
-0.750 -0.0430 0.01620 0.00561 -0.0061 1.0000 1.0000
-0.500 -0.0283 0.01610 0.00548 -0.0041 1.0000 1.0000
-0.250 -0.0141 0.01605 0.00541 -0.0021 1.0000 1.0000
0.000 0.0000 0.01603 0.00538 0.0000 1.0000 1.0000
0.250 0.0141 0.01605 0.00541 0.0021 1.0000 1.0000
0.500 0.0283 0.01610 0.00548 0.0041 1.0000 1.0000
0.750 0.0430 0.01620 0.00561 0.0061 1.0000 1.0000
1.000 0.0583 0.01634 0.00579 0.0078 1.0000 1.0000
1.250 0.0742 0.01653 0.00603 0.0094 1.0000 1.0000
1.500 0.0905 0.01677 0.00633 0.0107 1.0000 1.0000
1.750 0.1071 0.01707 0.00673 0.0119 1.0000 1.0000
2.000 0.1240 0.01743 0.00718 0.0128 1.0000 1.0000
2.250 0.1410 0.01786 0.00773 0.0136 1.0000 1.0000
2.500 0.1580 0.01837 0.00837 0.0142 1.0000 1.0000
2.750 0.1751 0.01895 0.00910 0.0145 1.0000 1.0000
3.000 0.2063 0.01967 0.01009 0.0121 0.9942 1.0000
3.250 0.3865 0.01785 0.00975 -0.0094 0.8523 1.0000
3.500 0.4036 0.01850 0.00829 0.0013 0.3938 1.0000
3.750 0.4188 0.02079 0.00930 0.0035 0.2669 1.0000
4.000 0.4445 0.02305 0.01115 0.0043 0.2169 1.0000
4.250 0.4731 0.02510 0.01313 0.0050 0.1913 1.0000
4.500 0.4998 0.02698 0.01500 0.0057 0.1732 1.0000
4.750 0.5271 0.02908 0.01728 0.0064 0.1650 1.0000
5.000 0.5540 0.03120 0.01975 0.0072 0.1598 1.0000
5.250 0.5789 0.03386 0.02246 0.0077 0.1536 1.0000
5.500 0.6030 0.03609 0.02538 0.0086 0.1488 1.0000
5.750 0.6264 0.03910 0.02890 0.0094 0.1494 1.0000
6.000 0.6484 0.04257 0.03281 0.0100 0.1517 1.0000
6.250 0.6682 0.04623 0.03690 0.0106 0.1526 1.0000
6.500 0.6868 0.05000 0.04118 0.0111 0.1557 1.0000
6.750 0.7003 0.05622 0.04829 0.0099 0.1808 1.0000
7.250 0.5827 0.08708 0.08005 -0.0419 0.5099 1.0000
7.500 0.6110 0.09211 0.08514 -0.0415 0.4890 1.0000
7.750 0.6183 0.09511 0.08813 -0.0394 0.4591 1.0000
8.000 0.6203 0.09767 0.09065 -0.0372 0.4279 1.0000
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Polar data table (+)
Polar graphs
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