MH 20 9.01% (mh20-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: MH 20 9.01% (mh20-il) Reynolds number: 500,000 Max Cl/Cd: 91.55 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-mh20-il-500000.txt Download as CSV file: xf-mh20-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5007 0.08336 0.08146 -0.0075 1.0000 0.0153
-8.000 -0.5029 0.07860 0.07674 -0.0113 1.0000 0.0153
-7.750 -0.5076 0.07368 0.07185 -0.0158 1.0000 0.0153
-7.500 -0.5055 0.06814 0.06629 -0.0200 1.0000 0.0153
-6.750 -0.5226 0.04752 0.04489 -0.0250 0.8727 0.0163
-6.500 -0.5168 0.04354 0.04067 -0.0245 0.8580 0.0168
-6.250 -0.5045 0.04086 0.03782 -0.0240 0.8457 0.0175
-6.000 -0.4905 0.03809 0.03482 -0.0233 0.8348 0.0182
-5.750 -0.4744 0.03518 0.03166 -0.0227 0.8245 0.0192
-5.500 -0.4569 0.03217 0.02830 -0.0217 0.8157 0.0206
-5.250 -0.4375 0.02942 0.02519 -0.0205 0.8079 0.0226
-5.000 -0.4083 0.03022 0.02568 -0.0194 0.7993 0.0254
-4.750 -0.3981 0.02363 0.01845 -0.0177 0.7930 0.0274
-4.500 -0.3735 0.01795 0.01217 -0.0149 0.7866 0.0120
-4.000 -0.3235 0.01370 0.00723 -0.0122 0.7741 0.0089
-3.750 -0.2981 0.01277 0.00612 -0.0114 0.7685 0.0085
-3.500 -0.2725 0.01197 0.00521 -0.0107 0.7623 0.0084
-3.250 -0.2470 0.01126 0.00436 -0.0099 0.7565 0.0085
-3.000 -0.2208 0.01072 0.00366 -0.0093 0.7507 0.0089
-2.750 -0.1939 0.01030 0.00307 -0.0088 0.7446 0.0105
-2.500 -0.1693 0.00954 0.00255 -0.0081 0.7393 0.0815
-2.250 -0.1443 0.00895 0.00236 -0.0078 0.7332 0.1799
-2.000 -0.1196 0.00846 0.00221 -0.0074 0.7281 0.2802
-1.750 -0.0982 0.00757 0.00208 -0.0065 0.7225 0.4792
-1.500 -0.0790 0.00680 0.00197 -0.0047 0.7171 0.6670
-1.250 -0.0585 0.00618 0.00193 -0.0027 0.7123 0.8364
-1.000 0.0155 0.00617 0.00205 -0.0121 0.7066 0.9458
-0.750 0.0531 0.00634 0.00210 -0.0139 0.7015 0.9694
-0.500 0.0968 0.00646 0.00214 -0.0172 0.6959 0.9825
-0.250 0.1451 0.00652 0.00209 -0.0215 0.6902 0.9913
0.000 0.1914 0.00653 0.00202 -0.0255 0.6846 0.9978
0.250 0.2253 0.00650 0.00194 -0.0269 0.6783 1.0000
0.500 0.2505 0.00652 0.00188 -0.0264 0.6730 1.0000
0.750 0.2761 0.00650 0.00185 -0.0260 0.6662 1.0000
1.000 0.3016 0.00652 0.00181 -0.0256 0.6604 1.0000
1.250 0.3276 0.00652 0.00180 -0.0252 0.6534 1.0000
1.500 0.3533 0.00654 0.00177 -0.0248 0.6470 1.0000
1.750 0.3794 0.00654 0.00179 -0.0245 0.6393 1.0000
2.000 0.4053 0.00658 0.00178 -0.0241 0.6321 1.0000
2.250 0.4315 0.00658 0.00180 -0.0237 0.6235 1.0000
2.500 0.4576 0.00661 0.00181 -0.0234 0.6147 1.0000
2.750 0.4836 0.00665 0.00183 -0.0230 0.6053 1.0000
3.000 0.5098 0.00668 0.00189 -0.0227 0.5941 1.0000
3.250 0.5359 0.00672 0.00194 -0.0223 0.5814 1.0000
3.500 0.5621 0.00677 0.00199 -0.0220 0.5671 1.0000
3.750 0.5882 0.00684 0.00205 -0.0217 0.5502 1.0000
4.000 0.6143 0.00693 0.00216 -0.0213 0.5294 1.0000
4.250 0.6401 0.00707 0.00225 -0.0210 0.5043 1.0000
4.500 0.6656 0.00727 0.00238 -0.0206 0.4758 1.0000
4.750 0.6908 0.00755 0.00256 -0.0202 0.4445 1.0000
5.000 0.7158 0.00786 0.00278 -0.0199 0.4111 1.0000
5.250 0.7404 0.00823 0.00304 -0.0195 0.3753 1.0000
5.500 0.7642 0.00872 0.00334 -0.0191 0.3258 1.0000
5.750 0.7871 0.00935 0.00374 -0.0186 0.2641 1.0000
6.000 0.8087 0.01015 0.00420 -0.0181 0.1941 1.0000
6.250 0.8293 0.01104 0.00476 -0.0174 0.1278 1.0000
6.500 0.8469 0.01234 0.00557 -0.0164 0.0486 1.0000
6.750 0.8631 0.01391 0.00697 -0.0147 0.0072 1.0000
7.000 0.8837 0.01469 0.00788 -0.0135 0.0061 1.0000
7.250 0.9031 0.01557 0.00888 -0.0123 0.0056 1.0000
7.500 0.9204 0.01663 0.01008 -0.0107 0.0053 1.0000
7.750 0.9357 0.01786 0.01144 -0.0088 0.0052 1.0000
8.000 0.9500 0.01915 0.01286 -0.0069 0.0052 1.0000
8.250 0.9628 0.02067 0.01452 -0.0047 0.0052 1.0000
8.500 0.9756 0.02240 0.01641 -0.0024 0.0054 1.0000
8.750 0.9882 0.02477 0.01900 -0.0001 0.0058 1.0000
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Polar data table (+)
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