MH 110 10.01% (mh110-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: MH 110 10.01% (mh110-il) Reynolds number: 50,000 Max Cl/Cd: 25.19 at α=9.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-mh110-il-50000-n5.txt Download as CSV file: xf-mh110-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: MH 110 10.01%
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 -0.6771 0.08131 0.07493 0.0017 1.0000 0.0300
-9.000 -0.6883 0.07661 0.07019 -0.0003 1.0000 0.0297
-8.750 -0.7017 0.07230 0.06581 -0.0009 1.0000 0.0296
-8.500 -0.7117 0.06806 0.06144 -0.0010 1.0000 0.0295
-8.250 -0.7181 0.06382 0.05699 -0.0007 1.0000 0.0293
-8.000 -0.7211 0.05960 0.05251 -0.0001 1.0000 0.0293
-7.750 -0.7206 0.05541 0.04798 0.0008 1.0000 0.0293
-7.500 -0.7162 0.05139 0.04356 0.0020 1.0000 0.0293
-7.250 -0.7077 0.04762 0.03934 0.0033 1.0000 0.0295
-7.000 -0.6955 0.04410 0.03531 0.0047 1.0000 0.0298
-6.750 -0.6817 0.04071 0.03164 0.0059 1.0000 0.0313
-6.500 -0.6644 0.03831 0.02905 0.0069 1.0000 0.0338
-6.250 -0.6443 0.03588 0.02620 0.0080 1.0000 0.0366
-6.000 -0.6211 0.03331 0.02320 0.0091 1.0000 0.0391
-5.750 -0.5955 0.03094 0.02045 0.0102 1.0000 0.0415
-5.500 -0.5733 0.02892 0.01846 0.0111 1.0000 0.0473
-5.250 -0.5494 0.02716 0.01655 0.0121 1.0000 0.0562
-5.000 -0.5264 0.02543 0.01478 0.0133 1.0000 0.0701
-4.750 -0.5040 0.02368 0.01318 0.0142 1.0000 0.0990
-4.500 -0.4826 0.02176 0.01177 0.0150 1.0000 0.1692
-4.250 -0.4637 0.02012 0.01081 0.0160 1.0000 0.2804
-4.000 -0.4469 0.01887 0.01012 0.0179 1.0000 0.3996
-3.750 -0.4298 0.01793 0.00972 0.0205 1.0000 0.5147
-3.500 -0.4078 0.01729 0.00952 0.0229 1.0000 0.6275
-3.250 -0.3751 0.01705 0.00949 0.0239 1.0000 0.7309
-3.000 -0.3309 0.01716 0.00951 0.0228 1.0000 0.8078
-2.750 -0.2846 0.01736 0.00942 0.0207 1.0000 0.8608
-2.500 -0.2358 0.01754 0.00934 0.0176 1.0000 0.9003
-2.250 -0.1809 0.01763 0.00917 0.0128 1.0000 0.9309
-2.000 -0.1269 0.01757 0.00888 0.0076 1.0000 0.9581
-1.750 -0.0567 0.01757 0.00854 -0.0011 0.8766 0.9821
-1.500 -0.0017 0.01753 0.00799 -0.0063 0.7912 1.0000
-1.250 0.0165 0.01760 0.00769 -0.0046 0.7419 1.0000
-1.000 0.0350 0.01768 0.00745 -0.0029 0.7037 1.0000
-0.750 0.0546 0.01775 0.00725 -0.0015 0.6707 1.0000
-0.500 0.0751 0.01783 0.00708 -0.0002 0.6419 1.0000
-0.250 0.0962 0.01793 0.00693 0.0009 0.6160 1.0000
0.000 0.1176 0.01806 0.00684 0.0020 0.5926 1.0000
0.250 0.1398 0.01819 0.00680 0.0029 0.5701 1.0000
0.500 0.1620 0.01836 0.00678 0.0039 0.5505 1.0000
0.750 0.1848 0.01853 0.00682 0.0047 0.5311 1.0000
1.000 0.2077 0.01873 0.00688 0.0055 0.5132 1.0000
1.250 0.2308 0.01895 0.00698 0.0063 0.4972 1.0000
1.500 0.2541 0.01918 0.00711 0.0071 0.4820 1.0000
1.750 0.2775 0.01944 0.00730 0.0078 0.4677 1.0000
2.000 0.3011 0.01972 0.00753 0.0085 0.4543 1.0000
2.250 0.3246 0.02003 0.00779 0.0092 0.4417 1.0000
2.500 0.3481 0.02036 0.00809 0.0099 0.4298 1.0000
2.750 0.3716 0.02072 0.00847 0.0106 0.4192 1.0000
3.000 0.3954 0.02109 0.00878 0.0113 0.4097 1.0000
3.250 0.4194 0.02151 0.00926 0.0118 0.3993 1.0000
3.500 0.4432 0.02197 0.00976 0.0124 0.3899 1.0000
3.750 0.4667 0.02241 0.01021 0.0132 0.3825 1.0000
4.000 0.4906 0.02297 0.01090 0.0137 0.3735 1.0000
4.250 0.5141 0.02348 0.01144 0.0144 0.3663 1.0000
4.500 0.5376 0.02408 0.01217 0.0149 0.3584 1.0000
4.750 0.5611 0.02469 0.01286 0.0155 0.3522 1.0000
5.000 0.5843 0.02540 0.01378 0.0160 0.3451 1.0000
5.250 0.6076 0.02603 0.01448 0.0168 0.3393 1.0000
5.500 0.6303 0.02687 0.01557 0.0172 0.3320 1.0000
5.750 0.6535 0.02752 0.01628 0.0180 0.3272 1.0000
6.000 0.6753 0.02861 0.01769 0.0183 0.3202 1.0000
6.250 0.6979 0.02935 0.01862 0.0191 0.3147 1.0000
6.500 0.7190 0.03048 0.02003 0.0195 0.3084 1.0000
6.750 0.7405 0.03148 0.02125 0.0202 0.3025 1.0000
7.000 0.7617 0.03248 0.02246 0.0209 0.2967 1.0000
7.250 0.7812 0.03372 0.02403 0.0215 0.2894 1.0000
7.500 0.8021 0.03464 0.02514 0.0223 0.2832 1.0000
7.750 0.8205 0.03586 0.02668 0.0230 0.2748 1.0000
8.000 0.8383 0.03702 0.02822 0.0238 0.2662 1.0000
8.250 0.8619 0.03700 0.02829 0.0252 0.2572 1.0000
8.500 0.8799 0.03755 0.02913 0.0263 0.2449 1.0000
8.750 0.9001 0.03728 0.02905 0.0277 0.2304 1.0000
9.000 0.9135 0.03772 0.02984 0.0289 0.2123 1.0000
9.250 0.9308 0.03695 0.02924 0.0304 0.1903 1.0000
9.500 0.9404 0.03753 0.03013 0.0314 0.1642 1.0000
9.750 0.9456 0.03880 0.03153 0.0322 0.1322 1.0000
10.000 0.9433 0.04098 0.03350 0.0329 0.1067 1.0000
10.250 0.9334 0.04418 0.03657 0.0333 0.0941 1.0000
10.500 0.9184 0.04786 0.04020 0.0329 0.0877 1.0000
10.750 0.9047 0.05229 0.04462 0.0312 0.0823 1.0000
11.000 0.8924 0.05695 0.04929 0.0291 0.0767 1.0000
11.250 0.8813 0.06149 0.05372 0.0272 0.0718 1.0000
11.500 0.8721 0.06610 0.05843 0.0253 0.0674 1.0000
11.750 0.8636 0.07046 0.06272 0.0236 0.0628 1.0000
12.000 0.8564 0.07487 0.06713 0.0222 0.0591 1.0000
12.250 0.8488 0.07967 0.07208 0.0205 0.0561 1.0000
12.500 0.8423 0.08425 0.07671 0.0188 0.0529 1.0000
12.750 0.8402 0.08769 0.08005 0.0179 0.0491 1.0000
13.000 0.8309 0.09338 0.08595 0.0154 0.0480 1.0000
13.250 0.8174 0.10013 0.09289 0.0122 0.0467 1.0000
13.500 0.8044 0.10713 0.09994 0.0089 0.0466 1.0000
13.750 0.7867 0.11574 0.10865 0.0047 0.0467 1.0000
14.000 0.7696 0.12479 0.11772 0.0003 0.0474 1.0000
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Polar data table (+)
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