MH 20 9.01% (mh20-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: MH 20 9.01% (mh20-il) Reynolds number: 1,000,000 Max Cl/Cd: 105.53 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-mh20-il-1000000.txt Download as CSV file: xf-mh20-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: MH 20 9.01%
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-3.750 -0.3078 0.01153 0.00573 -0.0089 0.7167 0.0062
-3.500 -0.2831 0.01040 0.00441 -0.0079 0.7113 0.0057
-3.250 -0.2574 0.00963 0.00348 -0.0072 0.7055 0.0054
-2.750 -0.2037 0.00879 0.00239 -0.0063 0.6950 0.0053
-2.500 -0.1763 0.00855 0.00203 -0.0060 0.6898 0.0055
-2.250 -0.1487 0.00837 0.00176 -0.0057 0.6850 0.0067
-2.000 -0.1231 0.00778 0.00147 -0.0053 0.6800 0.0829
-1.750 -0.0974 0.00735 0.00134 -0.0051 0.6753 0.1666
-1.250 -0.0475 0.00634 0.00114 -0.0043 0.6659 0.4053
-1.000 -0.0247 0.00574 0.00105 -0.0035 0.6613 0.5721
-0.750 -0.0014 0.00530 0.00100 -0.0026 0.6568 0.6903
-0.500 0.0190 0.00478 0.00096 -0.0008 0.6520 0.8316
-0.250 0.0707 0.00463 0.00107 -0.0057 0.6469 0.9449
0.250 0.1483 0.00484 0.00118 -0.0102 0.6362 0.9780
0.500 0.1871 0.00492 0.00122 -0.0125 0.6308 0.9839
0.750 0.2281 0.00496 0.00123 -0.0154 0.6245 0.9885
1.000 0.2657 0.00505 0.00126 -0.0176 0.6185 0.9928
1.250 0.3061 0.00506 0.00125 -0.0203 0.6112 0.9956
1.500 0.3472 0.00506 0.00122 -0.0233 0.6042 0.9988
1.750 0.3804 0.00506 0.00121 -0.0245 0.5960 1.0000
2.000 0.4065 0.00508 0.00120 -0.0242 0.5876 1.0000
2.250 0.4326 0.00511 0.00120 -0.0239 0.5777 1.0000
2.500 0.4588 0.00514 0.00121 -0.0236 0.5665 1.0000
2.750 0.4851 0.00519 0.00122 -0.0233 0.5529 1.0000
3.000 0.5113 0.00524 0.00127 -0.0230 0.5357 1.0000
3.250 0.5376 0.00534 0.00131 -0.0227 0.5143 1.0000
3.500 0.5637 0.00548 0.00137 -0.0225 0.4891 1.0000
3.750 0.5898 0.00566 0.00146 -0.0222 0.4614 1.0000
4.000 0.6158 0.00586 0.00159 -0.0220 0.4344 1.0000
4.250 0.6416 0.00608 0.00173 -0.0218 0.4070 1.0000
4.500 0.6673 0.00633 0.00188 -0.0215 0.3777 1.0000
4.750 0.6922 0.00674 0.00209 -0.0213 0.3265 1.0000
5.000 0.7163 0.00732 0.00236 -0.0210 0.2597 1.0000
5.250 0.7400 0.00793 0.00267 -0.0207 0.1964 1.0000
5.500 0.7640 0.00841 0.00298 -0.0203 0.1560 1.0000
5.750 0.7868 0.00909 0.00343 -0.0198 0.1023 1.0000
6.000 0.8081 0.00997 0.00398 -0.0191 0.0429 1.0000
6.250 0.8285 0.01099 0.00478 -0.0182 0.0039 1.0000
6.500 0.8521 0.01144 0.00532 -0.0174 0.0032 1.0000
6.750 0.8750 0.01195 0.00593 -0.0166 0.0030 1.0000
7.000 0.8970 0.01258 0.00666 -0.0157 0.0029 1.0000
7.250 0.9179 0.01331 0.00752 -0.0146 0.0029 1.0000
7.500 0.9368 0.01427 0.00860 -0.0132 0.0029 1.0000
7.750 0.9532 0.01544 0.00993 -0.0114 0.0029 1.0000
8.000 0.9680 0.01675 0.01138 -0.0095 0.0030 1.0000
8.250 0.9804 0.01839 0.01318 -0.0071 0.0032 1.0000
8.500 0.9926 0.02023 0.01517 -0.0048 0.0033 1.0000
8.750 1.0018 0.02314 0.01835 -0.0021 0.0036 1.0000
9.000 1.0194 0.02448 0.01974 -0.0006 0.0044 1.0000
12.000 0.9336 0.06658 0.06440 0.0172 0.0064 1.0000
12.250 0.9183 0.07196 0.06990 0.0143 0.0064 1.0000
12.500 0.8942 0.07973 0.07781 0.0096 0.0066 1.0000
12.750 0.8741 0.08762 0.08581 0.0043 0.0067 1.0000
13.000 0.8435 0.09927 0.09757 -0.0035 0.0073 1.0000
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Polar data table (+)
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