MH33 (mh33-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: MH33 (mh33-il) Reynolds number: 50,000 Max Cl/Cd: 32.3 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-mh33-il-50000.txt Download as CSV file: xf-mh33-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: MH33
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5223 0.10041 0.09369 0.0014 1.0000 0.2551
-8.000 -0.5283 0.09783 0.09122 0.0008 1.0000 0.2665
-7.500 -0.5320 0.09194 0.08550 0.0013 1.0000 0.2927
-7.250 -0.5323 0.08889 0.08254 0.0020 1.0000 0.3061
-6.500 -0.5115 0.07779 0.07161 0.0058 1.0000 0.3507
-5.750 -0.5189 0.05112 0.04361 -0.0294 1.0000 0.1155
-5.500 -0.5045 0.04599 0.03789 -0.0292 1.0000 0.1026
-5.250 -0.4881 0.04175 0.03335 -0.0282 1.0000 0.0968
-5.000 -0.4690 0.03785 0.02847 -0.0269 1.0000 0.0902
-4.750 -0.4483 0.03438 0.02462 -0.0256 1.0000 0.0890
-4.500 -0.4256 0.03136 0.02109 -0.0241 1.0000 0.0894
-4.250 -0.4030 0.02867 0.01819 -0.0228 1.0000 0.0962
-4.000 -0.3796 0.02651 0.01582 -0.0215 1.0000 0.1108
-3.750 -0.3531 0.02404 0.01311 -0.0197 1.0000 0.1310
-3.500 -0.3291 0.02158 0.01104 -0.0182 1.0000 0.1980
-3.250 -0.3142 0.01822 0.00941 -0.0153 1.0000 0.4263
-3.000 -0.1572 0.01671 0.00836 -0.0303 1.0000 1.0000
-2.750 -0.1522 0.01623 0.00762 -0.0278 1.0000 1.0000
-2.500 -0.1426 0.01593 0.00703 -0.0255 1.0000 1.0000
-2.250 -0.1304 0.01574 0.00656 -0.0233 1.0000 1.0000
-2.000 -0.1167 0.01560 0.00616 -0.0212 1.0000 1.0000
-1.750 -0.1021 0.01551 0.00583 -0.0192 1.0000 1.0000
-1.500 -0.0869 0.01546 0.00556 -0.0172 1.0000 1.0000
-1.250 -0.0709 0.01544 0.00528 -0.0154 1.0000 1.0000
-1.000 -0.0543 0.01545 0.00512 -0.0136 1.0000 1.0000
-0.750 -0.0371 0.01549 0.00500 -0.0120 1.0000 1.0000
-0.500 -0.0194 0.01555 0.00493 -0.0105 1.0000 1.0000
-0.250 -0.0013 0.01564 0.00490 -0.0090 1.0000 1.0000
0.000 0.0170 0.01576 0.00493 -0.0077 1.0000 1.0000
0.250 0.0354 0.01591 0.00496 -0.0064 1.0000 1.0000
0.500 0.0539 0.01609 0.00508 -0.0053 1.0000 1.0000
0.750 0.0723 0.01630 0.00526 -0.0042 1.0000 1.0000
1.000 0.0906 0.01655 0.00549 -0.0031 1.0000 1.0000
1.250 0.1086 0.01685 0.00579 -0.0021 1.0000 1.0000
1.500 0.1263 0.01720 0.00616 -0.0012 1.0000 1.0000
1.750 0.1434 0.01760 0.00660 -0.0004 1.0000 1.0000
2.000 0.1600 0.01808 0.00712 0.0004 1.0000 1.0000
2.250 0.1758 0.01864 0.00773 0.0010 1.0000 1.0000
2.500 0.1908 0.01930 0.00847 0.0015 1.0000 1.0000
2.750 0.2048 0.02008 0.00933 0.0018 1.0000 1.0000
3.000 0.2693 0.02152 0.01110 -0.0076 0.9746 1.0000
3.250 0.3611 0.02248 0.01252 -0.0204 0.9312 1.0000
3.500 0.4520 0.02229 0.01299 -0.0308 0.8820 1.0000
3.750 0.5334 0.02102 0.01234 -0.0368 0.8267 1.0000
4.000 0.5841 0.01983 0.01147 -0.0362 0.7629 1.0000
4.250 0.6145 0.01937 0.01109 -0.0325 0.6873 1.0000
4.500 0.6353 0.01967 0.01113 -0.0283 0.6021 1.0000
4.750 0.6533 0.02047 0.01154 -0.0245 0.5140 1.0000
5.000 0.6708 0.02164 0.01223 -0.0213 0.4299 1.0000
5.250 0.6883 0.02307 0.01329 -0.0186 0.3525 1.0000
5.500 0.7068 0.02469 0.01458 -0.0164 0.2865 1.0000
5.750 0.7275 0.02665 0.01633 -0.0145 0.2339 1.0000
6.000 0.7480 0.02868 0.01822 -0.0129 0.1915 1.0000
6.250 0.7700 0.03110 0.02050 -0.0115 0.1607 1.0000
6.500 0.7908 0.03362 0.02334 -0.0099 0.1382 1.0000
6.750 0.8118 0.03707 0.02741 -0.0083 0.1271 1.0000
7.000 0.8288 0.03987 0.03052 -0.0067 0.1130 1.0000
7.250 0.8445 0.04354 0.03454 -0.0051 0.1059 1.0000
7.500 0.8551 0.04802 0.03969 -0.0032 0.1044 1.0000
7.750 0.8612 0.05294 0.04519 -0.0016 0.1045 1.0000
8.000 0.8643 0.05813 0.05081 -0.0003 0.1058 1.0000
8.250 0.8681 0.06350 0.05644 0.0005 0.1075 1.0000
8.500 0.7802 0.06142 0.05537 0.0033 0.1151 1.0000
8.750 0.7629 0.06741 0.06146 0.0031 0.1183 1.0000
9.000 0.7717 0.07355 0.06763 0.0035 0.1217 1.0000
9.250 0.6979 0.08175 0.07579 -0.0039 0.1294 1.0000
9.500 0.7034 0.08778 0.08186 -0.0042 0.1353 1.0000
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