NASA SC(2)-0012 AIRFOIL (sc20012-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA SC(2)-0012 AIRFOIL (sc20012-il) Reynolds number: 50,000 Max Cl/Cd: 23.61 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-sc20012-il-50000-n5.txt Download as CSV file: xf-sc20012-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0012 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.7111 0.10060 0.09259 -0.0151 1.0000 0.0702
-11.250 -0.7442 0.08936 0.08131 -0.0228 1.0000 0.0694
-11.000 -0.7825 0.08028 0.07210 -0.0282 1.0000 0.0686
-10.750 -0.8150 0.07383 0.06547 -0.0301 1.0000 0.0682
-10.500 -0.8411 0.06907 0.06051 -0.0294 1.0000 0.0683
-10.250 -0.8601 0.06509 0.05629 -0.0276 1.0000 0.0688
-10.000 -0.8724 0.06121 0.05211 -0.0258 1.0000 0.0698
-9.750 -0.8807 0.05740 0.04789 -0.0239 1.0000 0.0712
-9.500 -0.8858 0.05356 0.04352 -0.0217 1.0000 0.0727
-9.250 -0.8784 0.05061 0.04030 -0.0202 1.0000 0.0744
-9.000 -0.8641 0.04852 0.03811 -0.0193 1.0000 0.0766
-8.750 -0.8509 0.04631 0.03568 -0.0181 1.0000 0.0795
-8.500 -0.8386 0.04372 0.03262 -0.0166 1.0000 0.0828
-8.250 -0.8220 0.04142 0.03003 -0.0155 1.0000 0.0857
-8.000 -0.8023 0.03982 0.02839 -0.0148 1.0000 0.0894
-7.750 -0.7827 0.03800 0.02629 -0.0138 1.0000 0.0938
-7.500 -0.7613 0.03620 0.02424 -0.0131 1.0000 0.0980
-7.250 -0.7391 0.03485 0.02289 -0.0124 1.0000 0.1028
-7.000 -0.7157 0.03341 0.02119 -0.0117 1.0000 0.1083
-6.750 -0.6918 0.03210 0.01993 -0.0112 1.0000 0.1138
-6.500 -0.6675 0.03096 0.01868 -0.0106 1.0000 0.1206
-6.250 -0.6429 0.02984 0.01758 -0.0100 1.0000 0.1274
-6.000 -0.6194 0.02889 0.01659 -0.0092 1.0000 0.1360
-5.750 -0.5972 0.02793 0.01569 -0.0082 1.0000 0.1446
-5.500 -0.5768 0.02706 0.01483 -0.0070 1.0000 0.1559
-5.250 -0.5586 0.02620 0.01404 -0.0054 1.0000 0.1691
-5.000 -0.5430 0.02533 0.01331 -0.0036 1.0000 0.1855
-4.750 -0.5290 0.02445 0.01260 -0.0016 1.0000 0.2086
-4.500 -0.5169 0.02350 0.01193 0.0007 1.0000 0.2424
-4.250 -0.5073 0.02241 0.01129 0.0032 1.0000 0.2980
-4.000 -0.5006 0.02121 0.01083 0.0064 1.0000 0.3998
-3.750 -0.4907 0.02078 0.01152 0.0112 1.0000 0.5637
-3.500 -0.4790 0.02103 0.01187 0.0152 1.0000 0.6649
-3.250 -0.4667 0.02127 0.01207 0.0189 1.0000 0.7183
-3.000 -0.4517 0.02160 0.01238 0.0224 1.0000 0.7587
-2.750 -0.4351 0.02194 0.01266 0.0256 1.0000 0.7921
-2.500 -0.4157 0.02230 0.01294 0.0283 1.0000 0.8220
-2.250 -0.3895 0.02282 0.01339 0.0301 1.0000 0.8524
-2.000 -0.3559 0.02341 0.01389 0.0305 1.0000 0.8844
-1.750 -0.3017 0.02407 0.01440 0.0268 1.0000 0.9107
-1.500 -0.2565 0.02432 0.01451 0.0234 1.0000 0.9273
-1.250 -0.2187 0.02439 0.01447 0.0208 1.0000 0.9398
-1.000 -0.1737 0.02442 0.01441 0.0166 0.9971 0.9489
-0.750 -0.1292 0.02442 0.01432 0.0123 0.9921 0.9565
-0.500 -0.0855 0.02440 0.01424 0.0081 0.9868 0.9629
-0.250 -0.0430 0.02442 0.01423 0.0041 0.9822 0.9702
0.000 0.0000 0.02440 0.01420 0.0000 0.9761 0.9761
0.250 0.0430 0.02441 0.01423 -0.0041 0.9702 0.9822
0.500 0.0855 0.02440 0.01424 -0.0081 0.9629 0.9868
0.750 0.1292 0.02442 0.01432 -0.0123 0.9565 0.9922
1.000 0.1737 0.02442 0.01440 -0.0166 0.9489 0.9971
1.250 0.2187 0.02439 0.01447 -0.0208 0.9398 1.0000
1.500 0.2564 0.02432 0.01450 -0.0234 0.9273 1.0000
1.750 0.3017 0.02407 0.01439 -0.0268 0.9107 1.0000
2.000 0.3558 0.02341 0.01388 -0.0305 0.8845 1.0000
2.250 0.3894 0.02282 0.01339 -0.0301 0.8524 1.0000
2.500 0.4157 0.02230 0.01293 -0.0283 0.8220 1.0000
2.750 0.4350 0.02193 0.01265 -0.0255 0.7922 1.0000
3.000 0.4516 0.02160 0.01238 -0.0224 0.7587 1.0000
3.250 0.4666 0.02126 0.01207 -0.0189 0.7184 1.0000
3.500 0.4789 0.02102 0.01187 -0.0152 0.6650 1.0000
3.750 0.4906 0.02078 0.01152 -0.0112 0.5640 1.0000
4.000 0.5005 0.02121 0.01083 -0.0064 0.4000 1.0000
4.250 0.5072 0.02241 0.01129 -0.0032 0.2981 1.0000
4.500 0.5168 0.02350 0.01192 -0.0007 0.2425 1.0000
4.750 0.5289 0.02445 0.01259 0.0016 0.2086 1.0000
5.000 0.5429 0.02532 0.01331 0.0036 0.1855 1.0000
5.250 0.5586 0.02620 0.01403 0.0055 0.1691 1.0000
5.500 0.5767 0.02706 0.01483 0.0070 0.1559 1.0000
5.750 0.5972 0.02793 0.01569 0.0082 0.1446 1.0000
6.000 0.6194 0.02889 0.01658 0.0092 0.1360 1.0000
6.250 0.6429 0.02984 0.01757 0.0100 0.1274 1.0000
6.500 0.6675 0.03096 0.01868 0.0106 0.1206 1.0000
6.750 0.6918 0.03210 0.01993 0.0112 0.1138 1.0000
7.000 0.7157 0.03341 0.02119 0.0117 0.1083 1.0000
7.250 0.7392 0.03485 0.02289 0.0124 0.1028 1.0000
7.500 0.7613 0.03620 0.02424 0.0130 0.0980 1.0000
7.750 0.7827 0.03800 0.02629 0.0138 0.0938 1.0000
8.000 0.8024 0.03982 0.02839 0.0148 0.0894 1.0000
8.250 0.8221 0.04142 0.03002 0.0155 0.0857 1.0000
8.500 0.8387 0.04372 0.03263 0.0166 0.0828 1.0000
8.750 0.8510 0.04631 0.03568 0.0180 0.0795 1.0000
9.000 0.8642 0.04852 0.03811 0.0192 0.0766 1.0000
9.250 0.8786 0.05061 0.04030 0.0202 0.0744 1.0000
9.500 0.8859 0.05357 0.04353 0.0216 0.0727 1.0000
9.750 0.8809 0.05741 0.04790 0.0238 0.0712 1.0000
10.000 0.8726 0.06123 0.05213 0.0258 0.0698 1.0000
10.250 0.8604 0.06511 0.05631 0.0276 0.0688 1.0000
10.500 0.8415 0.06910 0.06054 0.0293 0.0683 1.0000
10.750 0.8154 0.07386 0.06550 0.0300 0.0682 1.0000
11.000 0.7830 0.08033 0.07215 0.0280 0.0686 1.0000
11.250 0.7447 0.08945 0.08140 0.0226 0.0694 1.0000
11.500 0.7117 0.10073 0.09272 0.0148 0.0702 1.0000
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Polar data table (+)
Polar graphs
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