S5010 (s5010-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: S5010 (s5010-il) Reynolds number: 100,000 Max Cl/Cd: 45 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s5010-il-100000.txt Download as CSV file: xf-s5010-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: S5010
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4339 0.10562 0.10168 0.0148 1.0000 0.0958
-9.000 -0.4446 0.10182 0.09793 0.0107 1.0000 0.0993
-8.750 -0.4663 0.09817 0.09437 0.0046 1.0000 0.1004
-8.500 -0.4371 0.09243 0.08863 0.0083 1.0000 0.1038
-8.250 -0.4287 0.08854 0.08475 0.0078 1.0000 0.1085
-8.000 -0.4389 0.08429 0.08059 0.0037 1.0000 0.1130
-7.750 -0.5184 0.08764 0.08366 0.0063 1.0000 0.1059
-7.500 -0.5184 0.08389 0.07998 0.0033 1.0000 0.1101
-7.250 -0.5476 0.07926 0.07510 -0.0123 1.0000 0.1149
-7.000 -0.5172 0.07444 0.07059 -0.0059 1.0000 0.1187
-6.750 -0.5261 0.07097 0.06672 -0.0157 1.0000 0.1293
-6.500 -0.5012 0.06580 0.06189 -0.0124 1.0000 0.1328
-6.250 -0.4940 0.06165 0.05754 -0.0172 1.0000 0.1451
-6.000 -0.4799 0.05842 0.05419 -0.0195 1.0000 0.1590
-5.750 -0.4626 0.05497 0.05080 -0.0201 1.0000 0.1741
-5.500 -0.4466 0.05202 0.04797 -0.0197 1.0000 0.1903
-5.250 -0.4493 0.05035 0.04626 -0.0181 1.0000 0.2040
-5.000 -0.3428 0.01915 0.01331 -0.0349 0.9588 0.0835
-4.750 -0.3535 0.03112 0.02405 -0.0292 0.9696 0.0717
-4.500 -0.3140 0.02755 0.02004 -0.0316 0.9565 0.0699
-4.250 -0.2765 0.02551 0.01744 -0.0329 0.9426 0.0718
-4.000 -0.2434 0.02356 0.01509 -0.0332 0.9283 0.0720
-3.750 -0.2144 0.02198 0.01325 -0.0326 0.9140 0.0734
-3.500 -0.1895 0.02067 0.01199 -0.0318 0.8998 0.0793
-3.250 -0.1649 0.01986 0.01106 -0.0304 0.8860 0.0847
-3.000 -0.1425 0.01884 0.01010 -0.0288 0.8727 0.0907
-2.750 -0.1211 0.01810 0.00937 -0.0271 0.8598 0.1034
-2.500 -0.1008 0.01719 0.00866 -0.0253 0.8475 0.1395
-2.250 -0.0942 0.01408 0.00834 -0.0204 0.8372 0.7052
-2.000 0.0468 0.01392 0.00811 -0.0368 0.8302 1.0000
-1.750 0.0649 0.01395 0.00791 -0.0355 0.8169 1.0000
-1.500 0.0840 0.01402 0.00777 -0.0342 0.8042 1.0000
-1.250 0.1038 0.01413 0.00766 -0.0328 0.7922 1.0000
-1.000 0.1237 0.01425 0.00758 -0.0313 0.7810 1.0000
-0.750 0.1440 0.01438 0.00752 -0.0298 0.7699 1.0000
-0.500 0.1665 0.01453 0.00752 -0.0289 0.7578 1.0000
-0.250 0.1890 0.01470 0.00755 -0.0279 0.7463 1.0000
0.000 0.2110 0.01487 0.00757 -0.0266 0.7354 1.0000
0.250 0.2324 0.01500 0.00754 -0.0251 0.7256 1.0000
0.500 0.2562 0.01518 0.00762 -0.0243 0.7135 1.0000
0.750 0.2797 0.01537 0.00772 -0.0234 0.7019 1.0000
1.000 0.3029 0.01554 0.00778 -0.0223 0.6911 1.0000
1.250 0.3253 0.01564 0.00775 -0.0207 0.6816 1.0000
1.500 0.3497 0.01583 0.00789 -0.0200 0.6693 1.0000
1.750 0.3739 0.01602 0.00802 -0.0192 0.6575 1.0000
2.000 0.3978 0.01618 0.00812 -0.0182 0.6463 1.0000
2.250 0.4212 0.01627 0.00810 -0.0168 0.6364 1.0000
2.500 0.4456 0.01641 0.00820 -0.0159 0.6243 1.0000
2.750 0.4702 0.01659 0.00835 -0.0152 0.6117 1.0000
3.000 0.4947 0.01674 0.00848 -0.0143 0.5994 1.0000
3.250 0.5192 0.01688 0.00860 -0.0134 0.5873 1.0000
3.500 0.5438 0.01698 0.00865 -0.0123 0.5754 1.0000
3.750 0.5683 0.01704 0.00862 -0.0112 0.5639 1.0000
4.000 0.5932 0.01717 0.00876 -0.0104 0.5505 1.0000
4.250 0.6182 0.01731 0.00890 -0.0096 0.5366 1.0000
4.500 0.6433 0.01745 0.00903 -0.0088 0.5224 1.0000
4.750 0.6683 0.01759 0.00915 -0.0080 0.5080 1.0000
5.000 0.6934 0.01773 0.00930 -0.0072 0.4931 1.0000
5.250 0.7185 0.01788 0.00942 -0.0064 0.4780 1.0000
5.500 0.7435 0.01806 0.00959 -0.0056 0.4620 1.0000
5.750 0.7684 0.01832 0.00989 -0.0050 0.4445 1.0000
6.000 0.7932 0.01856 0.01018 -0.0043 0.4269 1.0000
6.250 0.8180 0.01882 0.01043 -0.0036 0.4095 1.0000
6.500 0.8427 0.01909 0.01065 -0.0029 0.3921 1.0000
6.750 0.8670 0.01947 0.01104 -0.0023 0.3737 1.0000
7.000 0.8910 0.01989 0.01154 -0.0017 0.3542 1.0000
7.250 0.9149 0.02033 0.01194 -0.0010 0.3357 1.0000
7.500 0.9383 0.02085 0.01244 -0.0004 0.3170 1.0000
7.750 0.9611 0.02147 0.01312 0.0002 0.2973 1.0000
8.000 0.9837 0.02209 0.01374 0.0009 0.2787 1.0000
8.250 1.0060 0.02283 0.01442 0.0015 0.2609 1.0000
8.500 1.0271 0.02366 0.01537 0.0021 0.2422 1.0000
8.750 1.0480 0.02454 0.01628 0.0028 0.2249 1.0000
9.000 1.0684 0.02548 0.01723 0.0035 0.2084 1.0000
9.250 1.0883 0.02649 0.01823 0.0043 0.1929 1.0000
9.500 1.1076 0.02762 0.01937 0.0050 0.1782 1.0000
9.750 1.1259 0.02882 0.02059 0.0058 0.1641 1.0000
10.000 1.1430 0.03011 0.02197 0.0067 0.1511 1.0000
10.250 1.1587 0.03148 0.02348 0.0076 0.1388 1.0000
10.500 1.1731 0.03304 0.02523 0.0087 0.1277 1.0000
10.750 1.1869 0.03473 0.02701 0.0097 0.1177 1.0000
11.000 1.2014 0.03636 0.02861 0.0106 0.1079 1.0000
11.250 1.2105 0.03795 0.03038 0.0118 0.0993 1.0000
11.500 1.2161 0.04009 0.03274 0.0132 0.0915 1.0000
11.750 1.2274 0.04202 0.03454 0.0141 0.0834 1.0000
12.000 1.2196 0.04405 0.03697 0.0161 0.0782 1.0000
12.250 1.2262 0.04610 0.03886 0.0172 0.0713 1.0000
12.500 1.2132 0.04873 0.04186 0.0185 0.0683 1.0000
12.750 1.2052 0.05135 0.04466 0.0189 0.0643 1.0000
13.000 1.2055 0.05424 0.04748 0.0193 0.0592 1.0000
13.250 1.1911 0.05815 0.05172 0.0187 0.0575 1.0000
13.500 1.1763 0.06246 0.05630 0.0176 0.0555 1.0000
13.750 1.1624 0.06706 0.06115 0.0161 0.0540 1.0000
14.000 1.1512 0.07148 0.06570 0.0144 0.0521 1.0000
14.250 0.8941 0.13260 0.12752 -0.0241 0.0953 1.0000
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Polar data table (+)
Polar graphs
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