NACA 64-012A AIRFOIL (n64012a-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 64-012A AIRFOIL (n64012a-il) Reynolds number: 100,000 Max Cl/Cd: 40 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012a-il-100000.txt Download as CSV file: xf-n64012a-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-012A 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.8266 0.05261 0.04601 -0.0215 1.0000 0.0784
-8.500 -0.8202 0.04834 0.04146 -0.0198 1.0000 0.0758
-8.250 -0.8137 0.04434 0.03702 -0.0175 1.0000 0.0734
-8.000 -0.8091 0.04047 0.03223 -0.0139 1.0000 0.0691
-7.750 -0.7919 0.03878 0.03006 -0.0117 1.0000 0.0679
-7.500 -0.7738 0.03586 0.02681 -0.0102 1.0000 0.0676
-7.250 -0.7546 0.03324 0.02388 -0.0088 1.0000 0.0679
-7.000 -0.7324 0.03007 0.02063 -0.0082 1.0000 0.0701
-6.750 -0.7109 0.02839 0.01891 -0.0072 1.0000 0.0738
-6.500 -0.6876 0.02674 0.01709 -0.0062 1.0000 0.0768
-6.250 -0.6640 0.02527 0.01543 -0.0050 1.0000 0.0799
-6.000 -0.6415 0.02347 0.01377 -0.0041 1.0000 0.0852
-5.750 -0.6221 0.02244 0.01271 -0.0025 1.0000 0.0933
-5.500 -0.6062 0.02105 0.01151 -0.0005 1.0000 0.1019
-5.250 -0.5934 0.01994 0.01051 0.0021 1.0000 0.1153
-5.000 -0.5831 0.01881 0.00954 0.0050 1.0000 0.1378
-4.750 -0.5800 0.01690 0.00834 0.0087 1.0000 0.2127
-4.500 -0.5890 0.01472 0.00819 0.0151 1.0000 0.5635
-4.250 -0.5766 0.01498 0.00860 0.0190 1.0000 0.6415
-4.000 -0.5615 0.01530 0.00891 0.0221 1.0000 0.6835
-3.750 -0.5455 0.01567 0.00925 0.0251 1.0000 0.7148
-3.500 -0.5296 0.01603 0.00958 0.0281 1.0000 0.7425
-3.250 -0.5134 0.01643 0.00999 0.0312 1.0000 0.7655
-3.000 -0.4999 0.01688 0.01043 0.0350 1.0000 0.7919
-2.750 -0.4885 0.01734 0.01087 0.0393 1.0000 0.8193
-2.500 -0.4749 0.01783 0.01136 0.0434 1.0000 0.8409
-2.250 -0.4427 0.01812 0.01155 0.0426 0.9954 0.8569
-2.000 -0.3997 0.01833 0.01161 0.0394 0.9875 0.8691
-1.750 -0.3567 0.01847 0.01162 0.0358 0.9796 0.8807
-1.500 -0.3082 0.01869 0.01173 0.0314 0.9726 0.8888
-1.250 -0.2643 0.01878 0.01172 0.0276 0.9646 0.8981
-1.000 -0.2131 0.01893 0.01178 0.0225 0.9581 0.9058
-0.750 -0.1653 0.01902 0.01181 0.0179 0.9508 0.9136
-0.500 -0.1110 0.01913 0.01187 0.0122 0.9446 0.9204
-0.250 -0.0546 0.01918 0.01187 0.0059 0.9398 0.9271
0.000 0.0000 0.01923 0.01192 0.0000 0.9326 0.9326
0.250 0.0546 0.01918 0.01187 -0.0059 0.9271 0.9398
0.500 0.1110 0.01913 0.01187 -0.0122 0.9204 0.9446
0.750 0.1653 0.01902 0.01180 -0.0179 0.9136 0.9509
1.000 0.2132 0.01893 0.01178 -0.0225 0.9058 0.9581
1.250 0.2643 0.01878 0.01171 -0.0276 0.8981 0.9646
1.500 0.3083 0.01869 0.01173 -0.0314 0.8888 0.9726
1.750 0.3566 0.01846 0.01161 -0.0358 0.8807 0.9797
2.000 0.3997 0.01833 0.01161 -0.0394 0.8691 0.9875
2.250 0.4426 0.01811 0.01154 -0.0426 0.8569 0.9954
2.500 0.4747 0.01782 0.01136 -0.0433 0.8409 1.0000
2.750 0.4883 0.01734 0.01087 -0.0393 0.8193 1.0000
3.000 0.4997 0.01688 0.01043 -0.0350 0.7919 1.0000
3.250 0.5132 0.01643 0.00999 -0.0312 0.7656 1.0000
3.500 0.5293 0.01602 0.00957 -0.0281 0.7425 1.0000
3.750 0.5452 0.01567 0.00925 -0.0251 0.7149 1.0000
4.000 0.5612 0.01530 0.00891 -0.0221 0.6836 1.0000
4.250 0.5764 0.01497 0.00860 -0.0190 0.6417 1.0000
4.500 0.5888 0.01472 0.00819 -0.0150 0.5641 1.0000
4.750 0.5800 0.01688 0.00833 -0.0086 0.2137 1.0000
5.000 0.5830 0.01881 0.00953 -0.0050 0.1379 1.0000
5.250 0.5933 0.01993 0.01050 -0.0021 0.1153 1.0000
5.500 0.6061 0.02104 0.01151 0.0005 0.1019 1.0000
5.750 0.6220 0.02244 0.01270 0.0025 0.0933 1.0000
6.000 0.6414 0.02347 0.01376 0.0041 0.0852 1.0000
6.250 0.6639 0.02527 0.01543 0.0050 0.0799 1.0000
6.500 0.6875 0.02674 0.01708 0.0062 0.0768 1.0000
6.750 0.7108 0.02839 0.01891 0.0073 0.0738 1.0000
7.000 0.7324 0.03007 0.02063 0.0082 0.0701 1.0000
7.250 0.7546 0.03325 0.02389 0.0088 0.0679 1.0000
7.500 0.7738 0.03586 0.02681 0.0103 0.0676 1.0000
7.750 0.7919 0.03878 0.03006 0.0117 0.0679 1.0000
8.000 0.8091 0.04048 0.03223 0.0139 0.0692 1.0000
8.250 0.8138 0.04434 0.03702 0.0175 0.0734 1.0000
8.500 0.8203 0.04835 0.04147 0.0197 0.0758 1.0000
8.750 0.8267 0.05263 0.04603 0.0214 0.0783 1.0000
9.000 0.7705 0.04873 0.04296 0.0273 0.0936 1.0000
9.500 0.6582 0.08988 0.08506 0.0073 0.1884 1.0000
9.750 0.7136 0.08920 0.08445 0.0169 0.1791 1.0000
10.000 0.6464 0.09910 0.09418 0.0037 0.1725 1.0000
10.250 0.6888 0.09968 0.09485 0.0114 0.1647 1.0000
10.500 0.6400 0.10783 0.10284 0.0007 0.1579 1.0000
10.750 0.5186 0.10436 0.09961 0.0126 0.1670 1.0000
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Polar data table (+)
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