MH 110 10.01% (mh110-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: MH 110 10.01% (mh110-il) Reynolds number: 500,000 Max Cl/Cd: 75.92 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-mh110-il-500000-n5.txt Download as CSV file: xf-mh110-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.7536 0.05935 0.05717 -0.0025 1.0000 0.0041
-9.250 -0.7793 0.05397 0.05164 -0.0027 1.0000 0.0040
-9.000 -0.8004 0.04952 0.04700 -0.0003 1.0000 0.0040
-8.750 -0.8166 0.04396 0.04113 0.0020 1.0000 0.0039
-8.500 -0.8256 0.03865 0.03542 0.0046 1.0000 0.0038
-8.250 -0.8272 0.03386 0.03018 0.0073 1.0000 0.0037
-8.000 -0.8221 0.02978 0.02564 0.0098 1.0000 0.0036
-7.750 -0.8114 0.02641 0.02182 0.0121 1.0000 0.0035
-7.500 -0.7961 0.02381 0.01884 0.0140 1.0000 0.0034
-7.250 -0.7777 0.02177 0.01650 0.0156 1.0000 0.0034
-7.000 -0.7576 0.02010 0.01459 0.0169 1.0000 0.0034
-6.750 -0.7364 0.01866 0.01295 0.0181 1.0000 0.0034
-6.500 -0.7145 0.01742 0.01155 0.0193 1.0000 0.0034
-6.250 -0.6921 0.01636 0.01035 0.0204 1.0000 0.0035
-6.000 -0.6694 0.01542 0.00930 0.0214 1.0000 0.0035
-5.750 -0.6464 0.01458 0.00834 0.0224 1.0000 0.0036
-5.500 -0.6209 0.01379 0.00747 0.0229 0.9798 0.0037
-5.250 -0.5867 0.01326 0.00661 0.0219 0.8450 0.0038
-5.000 -0.5654 0.01290 0.00598 0.0235 0.7852 0.0040
-4.750 -0.5420 0.01255 0.00539 0.0246 0.7411 0.0042
-4.500 -0.5176 0.01224 0.00488 0.0254 0.7039 0.0045
-4.250 -0.4926 0.01194 0.00439 0.0262 0.6709 0.0049
-4.000 -0.4676 0.01159 0.00389 0.0269 0.6411 0.0069
-3.750 -0.4420 0.01132 0.00348 0.0275 0.6139 0.0109
-3.500 -0.4170 0.01098 0.00314 0.0281 0.5898 0.0299
-3.250 -0.3931 0.01047 0.00281 0.0288 0.5666 0.0805
-2.750 -0.3449 0.00954 0.00232 0.0300 0.5254 0.2158
-2.500 -0.3196 0.00926 0.00212 0.0305 0.5064 0.2637
-2.250 -0.2944 0.00898 0.00195 0.0310 0.4886 0.3187
-2.000 -0.2693 0.00870 0.00181 0.0315 0.4718 0.3772
-1.750 -0.2444 0.00844 0.00168 0.0321 0.4556 0.4368
-1.500 -0.2194 0.00821 0.00157 0.0327 0.4404 0.4933
-1.250 -0.1949 0.00796 0.00147 0.0334 0.4262 0.5538
-1.000 -0.1708 0.00772 0.00139 0.0343 0.4136 0.6170
-0.750 -0.1468 0.00751 0.00134 0.0352 0.4014 0.6773
-0.500 -0.1226 0.00733 0.00131 0.0362 0.3890 0.7318
-0.250 -0.0977 0.00720 0.00128 0.0371 0.3766 0.7811
0.000 -0.0716 0.00711 0.00129 0.0377 0.3648 0.8256
0.250 -0.0428 0.00709 0.00133 0.0378 0.3539 0.8641
0.500 -0.0097 0.00714 0.00138 0.0370 0.3430 0.8949
0.750 0.0271 0.00724 0.00146 0.0353 0.3324 0.9165
1.000 0.0647 0.00738 0.00153 0.0333 0.3217 0.9311
1.250 0.0994 0.00752 0.00160 0.0320 0.3113 0.9424
1.500 0.1355 0.00765 0.00168 0.0303 0.3013 0.9499
1.750 0.1676 0.00778 0.00176 0.0296 0.2935 0.9573
2.250 0.2312 0.00804 0.00193 0.0281 0.2789 0.9690
2.500 0.2665 0.00818 0.00201 0.0265 0.2714 0.9720
2.750 0.3003 0.00829 0.00209 0.0253 0.2644 0.9758
3.250 0.3668 0.00854 0.00231 0.0230 0.2525 0.9837
3.500 0.4029 0.00866 0.00240 0.0212 0.2468 0.9866
3.750 0.4367 0.00878 0.00252 0.0199 0.2421 0.9896
4.000 0.4691 0.00888 0.00264 0.0189 0.2374 0.9922
4.250 0.5015 0.00901 0.00276 0.0179 0.2325 0.9943
4.500 0.5351 0.00912 0.00287 0.0166 0.2282 0.9961
4.750 0.5683 0.00921 0.00301 0.0154 0.2244 0.9980
5.000 0.6012 0.00934 0.00316 0.0142 0.2205 0.9997
5.250 0.6274 0.00950 0.00332 0.0144 0.2169 1.0000
5.500 0.6522 0.00962 0.00349 0.0150 0.2137 1.0000
5.750 0.6770 0.00975 0.00367 0.0156 0.2102 1.0000
6.000 0.7016 0.00992 0.00387 0.0162 0.2046 1.0000
6.250 0.7264 0.01006 0.00405 0.0168 0.1978 1.0000
6.500 0.7509 0.01026 0.00424 0.0173 0.1887 1.0000
6.750 0.7755 0.01043 0.00442 0.0179 0.1779 1.0000
7.000 0.8001 0.01062 0.00466 0.0185 0.1704 1.0000
7.250 0.8245 0.01086 0.00491 0.0191 0.1604 1.0000
7.500 0.8482 0.01119 0.00522 0.0197 0.1387 1.0000
7.750 0.8687 0.01209 0.00584 0.0204 0.0889 1.0000
8.000 0.8881 0.01314 0.00669 0.0213 0.0513 1.0000
8.250 0.9079 0.01406 0.00749 0.0221 0.0284 1.0000
8.500 0.9284 0.01483 0.00822 0.0230 0.0160 1.0000
8.750 0.9494 0.01548 0.00888 0.0238 0.0108 1.0000
9.000 0.9699 0.01617 0.00961 0.0247 0.0070 1.0000
9.250 0.9903 0.01686 0.01035 0.0256 0.0050 1.0000
9.500 1.0096 0.01766 0.01127 0.0266 0.0038 1.0000
9.750 1.0287 0.01844 0.01215 0.0277 0.0030 1.0000
10.000 1.0470 0.01926 0.01308 0.0287 0.0027 1.0000
10.250 1.0625 0.02037 0.01432 0.0300 0.0023 1.0000
10.500 1.0768 0.02153 0.01562 0.0313 0.0022 1.0000
10.750 1.0906 0.02265 0.01688 0.0326 0.0020 1.0000
11.000 1.1022 0.02388 0.01824 0.0340 0.0020 1.0000
11.250 1.1114 0.02523 0.01972 0.0355 0.0019 1.0000
11.500 1.1172 0.02672 0.02134 0.0371 0.0019 1.0000
11.750 1.1161 0.02843 0.02319 0.0391 0.0019 1.0000
12.000 1.1049 0.03107 0.02597 0.0401 0.0018 1.0000
12.250 1.0970 0.03513 0.03019 0.0375 0.0019 1.0000
12.500 1.0891 0.03984 0.03505 0.0345 0.0018 1.0000
12.750 1.0794 0.04477 0.04011 0.0316 0.0018 1.0000
13.000 1.0708 0.04942 0.04487 0.0292 0.0018 1.0000
13.250 1.0541 0.05514 0.05071 0.0263 0.0018 1.0000
13.500 1.0459 0.05964 0.05533 0.0243 0.0018 1.0000
13.750 1.0293 0.06534 0.06111 0.0217 0.0017 1.0000
14.000 1.0198 0.07006 0.06594 0.0195 0.0018 1.0000
14.250 1.0075 0.07520 0.07116 0.0172 0.0017 1.0000
14.500 0.9974 0.08023 0.07628 0.0150 0.0018 1.0000
14.750 0.9887 0.08521 0.08136 0.0127 0.0018 1.0000
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Polar data table (+)
Polar graphs
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