MH 110 10.01% (mh110-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: MH 110 10.01% (mh110-il) Reynolds number: 1,000,000 Max Cl/Cd: 88.89 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-mh110-il-1000000-n5.txt Download as CSV file: xf-mh110-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7714 0.06536 0.06393 0.0060 1.0000 0.0025
-10.000 -0.8092 0.05268 0.05100 -0.0042 1.0000 0.0025
-9.750 -0.8404 0.04753 0.04567 -0.0035 1.0000 0.0025
-9.500 -0.8605 0.04405 0.04201 -0.0001 1.0000 0.0025
-9.250 -0.8790 0.03891 0.03654 0.0028 1.0000 0.0025
-9.000 -0.8860 0.03450 0.03178 0.0057 1.0000 0.0024
-8.750 -0.8873 0.03031 0.02718 0.0085 1.0000 0.0024
-8.500 -0.8803 0.02708 0.02360 0.0109 1.0000 0.0024
-8.250 -0.8685 0.02441 0.02060 0.0129 1.0000 0.0023
-8.000 -0.8532 0.02224 0.01815 0.0147 1.0000 0.0023
-7.750 -0.8357 0.02033 0.01595 0.0163 1.0000 0.0023
-7.500 -0.8164 0.01873 0.01414 0.0177 1.0000 0.0023
-7.250 -0.7955 0.01742 0.01266 0.0189 1.0000 0.0023
-7.000 -0.7738 0.01627 0.01135 0.0201 1.0000 0.0023
-6.750 -0.7519 0.01518 0.01011 0.0212 1.0000 0.0023
-6.500 -0.7216 0.01434 0.00894 0.0208 0.8687 0.0023
-6.250 -0.7015 0.01381 0.00809 0.0225 0.7983 0.0023
-6.000 -0.6789 0.01324 0.00728 0.0236 0.7514 0.0023
-5.750 -0.6554 0.01271 0.00654 0.0246 0.7117 0.0023
-5.500 -0.6313 0.01222 0.00586 0.0254 0.6793 0.0023
-5.250 -0.6066 0.01179 0.00527 0.0261 0.6497 0.0024
-5.000 -0.5814 0.01142 0.00474 0.0268 0.6215 0.0024
-4.750 -0.5559 0.01108 0.00427 0.0273 0.5964 0.0025
-4.500 -0.5301 0.01081 0.00388 0.0278 0.5715 0.0025
-4.250 -0.5039 0.01057 0.00352 0.0283 0.5504 0.0026
-4.000 -0.4779 0.01029 0.00312 0.0288 0.5298 0.0030
-3.750 -0.4514 0.01009 0.00282 0.0292 0.5104 0.0035
-3.500 -0.4247 0.00992 0.00257 0.0296 0.4916 0.0049
-3.250 -0.3983 0.00971 0.00233 0.0300 0.4736 0.0111
-3.000 -0.3722 0.00947 0.00212 0.0304 0.4577 0.0274
-2.750 -0.3463 0.00919 0.00191 0.0308 0.4422 0.0568
-2.500 -0.3215 0.00877 0.00170 0.0313 0.4276 0.1187
-2.250 -0.2970 0.00832 0.00152 0.0319 0.4141 0.1962
-2.000 -0.2709 0.00810 0.00139 0.0322 0.4022 0.2390
-1.750 -0.2448 0.00789 0.00129 0.0326 0.3908 0.2839
-1.500 -0.2187 0.00768 0.00119 0.0330 0.3785 0.3311
-1.250 -0.1926 0.00750 0.00110 0.0333 0.3666 0.3785
-0.750 -0.1403 0.00715 0.00098 0.0341 0.3449 0.4757
-0.500 -0.1143 0.00698 0.00093 0.0345 0.3358 0.5250
-0.250 -0.0887 0.00682 0.00089 0.0350 0.3266 0.5779
0.000 -0.0631 0.00665 0.00086 0.0356 0.3181 0.6300
0.250 -0.0377 0.00651 0.00085 0.0362 0.3094 0.6792
0.500 -0.0124 0.00639 0.00084 0.0368 0.2999 0.7253
0.750 0.0128 0.00627 0.00085 0.0375 0.2921 0.7704
1.000 0.0380 0.00618 0.00087 0.0383 0.2848 0.8130
1.250 0.0638 0.00610 0.00090 0.0389 0.2783 0.8526
1.500 0.0911 0.00608 0.00096 0.0392 0.2713 0.8883
1.750 0.1220 0.00610 0.00103 0.0388 0.2653 0.9168
2.000 0.1570 0.00620 0.00113 0.0374 0.2578 0.9356
2.250 0.1926 0.00630 0.00122 0.0358 0.2512 0.9467
2.500 0.2250 0.00643 0.00130 0.0349 0.2448 0.9557
2.750 0.2600 0.00653 0.00138 0.0334 0.2402 0.9606
3.000 0.2882 0.00664 0.00147 0.0335 0.2354 0.9674
3.250 0.3211 0.00677 0.00157 0.0324 0.2304 0.9704
3.500 0.3545 0.00686 0.00167 0.0313 0.2267 0.9729
3.750 0.3867 0.00698 0.00176 0.0304 0.2222 0.9761
4.000 0.4151 0.00712 0.00190 0.0303 0.2173 0.9811
4.250 0.4501 0.00723 0.00201 0.0288 0.2142 0.9831
4.500 0.4846 0.00733 0.00211 0.0274 0.2108 0.9846
4.750 0.5174 0.00744 0.00222 0.0263 0.2073 0.9862
5.000 0.5494 0.00757 0.00236 0.0253 0.2031 0.9879
5.250 0.5810 0.00767 0.00248 0.0245 0.2005 0.9896
5.500 0.6119 0.00778 0.00261 0.0239 0.1971 0.9914
5.750 0.6419 0.00795 0.00276 0.0233 0.1899 0.9930
6.000 0.6731 0.00807 0.00290 0.0226 0.1828 0.9942
6.250 0.7052 0.00824 0.00304 0.0215 0.1713 0.9952
6.500 0.7369 0.00842 0.00321 0.0205 0.1621 0.9963
6.750 0.7680 0.00864 0.00339 0.0197 0.1488 0.9974
7.000 0.7966 0.00939 0.00386 0.0188 0.0960 0.9985
7.250 0.8244 0.01040 0.00459 0.0180 0.0460 0.9996
7.500 0.8502 0.01104 0.00514 0.0180 0.0251 1.0000
7.750 0.8730 0.01149 0.00556 0.0187 0.0168 1.0000
8.000 0.8958 0.01192 0.00599 0.0194 0.0119 1.0000
8.250 0.9186 0.01234 0.00641 0.0202 0.0084 1.0000
8.500 0.9408 0.01284 0.00692 0.0210 0.0050 1.0000
8.750 0.9631 0.01330 0.00741 0.0218 0.0033 1.0000
9.000 0.9848 0.01384 0.00798 0.0226 0.0021 1.0000
9.250 1.0067 0.01432 0.00853 0.0234 0.0017 1.0000
9.500 1.0280 0.01487 0.00915 0.0243 0.0014 1.0000
9.750 1.0483 0.01552 0.00988 0.0253 0.0012 1.0000
10.000 1.0686 0.01614 0.01058 0.0263 0.0011 1.0000
10.250 1.0867 0.01701 0.01156 0.0274 0.0010 1.0000
10.500 1.1057 0.01771 0.01234 0.0285 0.0010 1.0000
10.750 1.1232 0.01853 0.01326 0.0297 0.0010 1.0000
11.000 1.1399 0.01939 0.01421 0.0309 0.0010 1.0000
11.250 1.1555 0.02030 0.01522 0.0321 0.0010 1.0000
11.500 1.1695 0.02130 0.01633 0.0335 0.0010 1.0000
11.750 1.1815 0.02240 0.01755 0.0350 0.0010 1.0000
12.000 1.1901 0.02370 0.01897 0.0366 0.0009 1.0000
12.250 1.1960 0.02509 0.02048 0.0383 0.0009 1.0000
12.500 1.1921 0.02675 0.02226 0.0410 0.0009 1.0000
12.750 1.1837 0.02921 0.02485 0.0416 0.0009 1.0000
13.000 1.1791 0.03298 0.02875 0.0390 0.0009 1.0000
13.250 1.1662 0.03870 0.03465 0.0349 0.0009 1.0000
13.500 1.1608 0.04315 0.03920 0.0321 0.0009 1.0000
13.750 1.1395 0.04979 0.04600 0.0285 0.0009 1.0000
14.000 1.1290 0.05469 0.05098 0.0260 0.0009 1.0000
14.250 1.1098 0.06081 0.05722 0.0231 0.0009 1.0000
14.500 1.0950 0.06630 0.06279 0.0205 0.0009 1.0000
14.750 1.0812 0.07170 0.06828 0.0180 0.0009 1.0000
15.000 1.0622 0.07798 0.07466 0.0151 0.0009 1.0000
15.250 1.0528 0.08305 0.07980 0.0127 0.0009 1.0000
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Polar data table (+)
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