S5010 (s5010-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: S5010 (s5010-il) Reynolds number: 500,000 Max Cl/Cd: 80.77 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-s5010-il-500000-n5.txt Download as CSV file: xf-s5010-il-500000-n5.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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5824 0.08303 0.08077 0.0117 0.8365 0.0064
-8.500 -0.5952 0.07725 0.07497 0.0082 0.8163 0.0064
-8.250 -0.6150 0.07110 0.06877 0.0034 0.7988 0.0062
-8.000 -0.6295 0.06230 0.05983 -0.0028 0.7863 0.0062
-7.750 -0.6430 0.05199 0.04923 -0.0068 0.7764 0.0062
-7.500 -0.6832 0.02944 0.02531 -0.0068 0.7714 0.0066
-7.250 -0.6742 0.02310 0.01807 -0.0051 0.7627 0.0072
-7.000 -0.6539 0.02057 0.01505 -0.0041 0.7537 0.0076
-6.750 -0.6308 0.01896 0.01310 -0.0034 0.7446 0.0079
-6.500 -0.6060 0.01818 0.01218 -0.0030 0.7358 0.0084
-6.250 -0.5798 0.01789 0.01182 -0.0028 0.7272 0.0088
-6.000 -0.5535 0.01746 0.01127 -0.0025 0.7186 0.0093
-5.750 -0.5277 0.01673 0.01036 -0.0021 0.7107 0.0101
-5.500 -0.5018 0.01585 0.00929 -0.0016 0.7025 0.0108
-5.250 -0.4754 0.01521 0.00847 -0.0012 0.6948 0.0113
-5.000 -0.4499 0.01430 0.00743 -0.0007 0.6869 0.0120
-4.750 -0.4234 0.01382 0.00689 -0.0004 0.6793 0.0127
-4.500 -0.3965 0.01346 0.00647 -0.0002 0.6713 0.0136
-4.250 -0.3694 0.01311 0.00603 0.0000 0.6637 0.0148
-4.000 -0.3424 0.01273 0.00555 0.0003 0.6558 0.0156
-3.750 -0.3152 0.01238 0.00512 0.0006 0.6481 0.0162
-3.500 -0.2893 0.01168 0.00433 0.0010 0.6404 0.0175
-3.250 -0.2621 0.01134 0.00396 0.0013 0.6325 0.0188
-2.750 -0.2073 0.01076 0.00325 0.0017 0.6167 0.0207
-2.500 -0.1798 0.01053 0.00293 0.0019 0.6090 0.0216
-2.250 -0.1520 0.01032 0.00266 0.0021 0.6008 0.0227
-2.000 -0.1244 0.01008 0.00237 0.0023 0.5931 0.0253
-1.750 -0.0965 0.00991 0.00216 0.0024 0.5846 0.0280
-1.500 -0.0686 0.00974 0.00197 0.0026 0.5767 0.0364
-1.250 -0.0413 0.00947 0.00181 0.0027 0.5682 0.0745
-1.000 -0.0143 0.00911 0.00167 0.0028 0.5597 0.1426
-0.750 0.0119 0.00864 0.00154 0.0030 0.5513 0.2545
-0.500 0.0378 0.00816 0.00145 0.0032 0.5423 0.3806
-0.250 0.0633 0.00771 0.00138 0.0035 0.5338 0.5042
0.000 0.0875 0.00724 0.00135 0.0043 0.5248 0.6381
0.250 0.1078 0.00662 0.00135 0.0062 0.5163 0.8118
0.500 0.1595 0.00647 0.00148 0.0019 0.5052 0.9594
0.750 0.1972 0.00656 0.00148 0.0000 0.4945 0.9759
1.000 0.2326 0.00663 0.00147 -0.0015 0.4836 0.9850
1.250 0.2694 0.00670 0.00146 -0.0034 0.4725 0.9910
1.500 0.3058 0.00677 0.00146 -0.0052 0.4612 0.9960
1.750 0.3427 0.00684 0.00145 -0.0071 0.4492 1.0000
2.000 0.3689 0.00693 0.00147 -0.0068 0.4381 1.0000
2.250 0.3953 0.00702 0.00151 -0.0065 0.4268 1.0000
2.500 0.4216 0.00712 0.00155 -0.0061 0.4153 1.0000
2.750 0.4480 0.00723 0.00160 -0.0058 0.4034 1.0000
3.000 0.4744 0.00735 0.00168 -0.0055 0.3910 1.0000
3.250 0.5008 0.00749 0.00175 -0.0052 0.3779 1.0000
3.500 0.5272 0.00764 0.00185 -0.0050 0.3640 1.0000
3.750 0.5536 0.00780 0.00195 -0.0047 0.3496 1.0000
4.000 0.5799 0.00797 0.00208 -0.0044 0.3357 1.0000
4.250 0.6062 0.00815 0.00221 -0.0042 0.3212 1.0000
4.500 0.6325 0.00835 0.00235 -0.0039 0.3068 1.0000
4.750 0.6587 0.00856 0.00252 -0.0037 0.2918 1.0000
5.000 0.6849 0.00878 0.00270 -0.0035 0.2775 1.0000
5.250 0.7110 0.00901 0.00289 -0.0032 0.2632 1.0000
5.500 0.7370 0.00926 0.00309 -0.0030 0.2485 1.0000
5.750 0.7629 0.00953 0.00333 -0.0028 0.2346 1.0000
6.000 0.7887 0.00981 0.00357 -0.0025 0.2201 1.0000
6.250 0.8144 0.01009 0.00382 -0.0023 0.2069 1.0000
6.500 0.8400 0.01040 0.00409 -0.0021 0.1927 1.0000
6.750 0.8654 0.01073 0.00440 -0.0018 0.1793 1.0000
7.000 0.8907 0.01107 0.00471 -0.0016 0.1659 1.0000
7.250 0.9159 0.01143 0.00504 -0.0014 0.1541 1.0000
7.500 0.9411 0.01178 0.00537 -0.0011 0.1427 1.0000
7.750 0.9661 0.01216 0.00574 -0.0009 0.1307 1.0000
8.000 0.9908 0.01257 0.00614 -0.0006 0.1194 1.0000
8.250 1.0152 0.01300 0.00656 -0.0004 0.1086 1.0000
8.500 1.0394 0.01346 0.00701 -0.0001 0.0992 1.0000
8.750 1.0632 0.01395 0.00748 0.0002 0.0890 1.0000
9.000 1.0868 0.01444 0.00797 0.0005 0.0792 1.0000
9.250 1.1099 0.01499 0.00852 0.0008 0.0694 1.0000
9.500 1.1324 0.01558 0.00909 0.0012 0.0602 1.0000
9.750 1.1542 0.01622 0.00971 0.0016 0.0515 1.0000
10.000 1.1751 0.01692 0.01039 0.0021 0.0430 1.0000
10.250 1.1961 0.01758 0.01106 0.0025 0.0365 1.0000
10.500 1.2159 0.01831 0.01182 0.0031 0.0306 1.0000
10.750 1.2348 0.01909 0.01260 0.0037 0.0257 1.0000
11.000 1.2533 0.01984 0.01341 0.0044 0.0220 1.0000
11.250 1.2694 0.02076 0.01434 0.0052 0.0179 1.0000
11.500 1.2839 0.02172 0.01534 0.0062 0.0141 1.0000
11.750 1.2955 0.02282 0.01648 0.0074 0.0108 1.0000
12.000 1.3004 0.02407 0.01779 0.0091 0.0085 1.0000
12.250 1.3017 0.02564 0.01942 0.0107 0.0069 1.0000
12.500 1.3060 0.02727 0.02113 0.0115 0.0061 1.0000
12.750 1.3101 0.02908 0.02305 0.0121 0.0055 1.0000
13.000 1.3109 0.03135 0.02541 0.0123 0.0047 1.0000
13.250 1.3129 0.03361 0.02777 0.0124 0.0044 1.0000
13.500 1.3148 0.03600 0.03028 0.0122 0.0042 1.0000
13.750 1.3155 0.03861 0.03300 0.0119 0.0039 1.0000
14.000 1.3141 0.04155 0.03605 0.0113 0.0038 1.0000
14.250 1.3112 0.04476 0.03937 0.0105 0.0035 1.0000
14.500 1.3076 0.04817 0.04291 0.0095 0.0035 1.0000
14.750 1.3010 0.05206 0.04691 0.0083 0.0033 1.0000
15.000 1.2941 0.05614 0.05112 0.0068 0.0033 1.0000
15.250 1.2841 0.06076 0.05586 0.0050 0.0032 1.0000
15.500 1.2724 0.06594 0.06117 0.0028 0.0031 1.0000
15.750 1.2642 0.07076 0.06612 0.0006 0.0032 1.0000
16.000 1.2528 0.07628 0.07177 -0.0019 0.0031 1.0000
16.250 1.2367 0.08274 0.07836 -0.0051 0.0030 1.0000
16.500 1.2222 0.08917 0.08492 -0.0082 0.0030 1.0000
16.750 1.2064 0.09602 0.09190 -0.0116 0.0030 1.0000
17.000 1.1936 0.10246 0.09847 -0.0149 0.0030 1.0000
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Polar data table (+)
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