NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il) Reynolds number: 100,000 Max Cl/Cd: 39.12 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc410-il-100000.txt Download as CSV file: xf-rc410-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(4)-10 AIRFOIL
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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5157 0.10604 0.10135 0.0066 1.0000 0.1057
-8.750 -0.5386 0.10234 0.09776 -0.0042 1.0000 0.1078
-8.500 -0.5628 0.09845 0.09380 -0.0140 1.0000 0.1083
-8.250 -0.5160 0.09313 0.08862 -0.0025 1.0000 0.1111
-8.000 -0.5046 0.08980 0.08531 -0.0022 1.0000 0.1141
-7.750 -0.5034 0.08578 0.08135 -0.0058 1.0000 0.1173
-7.500 -0.5199 0.08121 0.07665 -0.0174 1.0000 0.1225
-7.250 -0.5200 0.07602 0.07136 -0.0211 1.0000 0.1241
-7.000 -0.5011 0.07156 0.06703 -0.0199 1.0000 0.1258
-6.750 -0.4866 0.06808 0.06359 -0.0201 1.0000 0.1285
-6.500 -0.4754 0.06458 0.06005 -0.0218 1.0000 0.1327
-6.250 -0.4867 0.06287 0.05782 -0.0241 1.0000 0.1394
-6.000 -0.4881 0.05943 0.05453 -0.0212 1.0000 0.1404
-5.750 -0.4901 0.05706 0.05223 -0.0181 1.0000 0.1417
-5.500 -0.4904 0.05495 0.05012 -0.0154 1.0000 0.1436
-5.250 -0.4629 0.04351 0.03740 -0.0200 0.9922 0.0960
-5.000 -0.4308 0.03909 0.03259 -0.0222 0.9856 0.0914
-4.750 -0.3991 0.03482 0.02781 -0.0241 0.9793 0.0900
-4.500 -0.3657 0.03124 0.02350 -0.0254 0.9730 0.0930
-4.250 -0.3301 0.02948 0.02166 -0.0278 0.9680 0.0997
-4.000 -0.2969 0.02761 0.01912 -0.0285 0.9614 0.1081
-3.750 -0.2578 0.02622 0.01770 -0.0312 0.9569 0.1206
-3.500 -0.2250 0.02512 0.01649 -0.0326 0.9507 0.1340
-3.250 -0.1912 0.02412 0.01534 -0.0340 0.9448 0.1486
-3.000 -0.1532 0.02310 0.01417 -0.0361 0.9407 0.1630
-2.750 -0.1253 0.02238 0.01331 -0.0363 0.9340 0.1742
-2.500 -0.0913 0.02187 0.01264 -0.0375 0.9284 0.1861
-2.250 -0.0535 0.02094 0.01181 -0.0395 0.9244 0.1976
-2.000 -0.0291 0.02050 0.01138 -0.0391 0.9174 0.2068
-1.750 0.0028 0.02012 0.01099 -0.0399 0.9123 0.2178
-1.500 0.0355 0.01956 0.01061 -0.0408 0.9074 0.2307
-1.250 0.0604 0.01921 0.01038 -0.0401 0.8981 0.2425
-1.000 0.0889 0.01882 0.01009 -0.0396 0.8883 0.2607
-0.750 0.1169 0.01833 0.00974 -0.0387 0.8772 0.2879
-0.500 0.2441 0.01564 0.00929 -0.0532 0.8664 1.0000
-0.250 0.2640 0.01579 0.00927 -0.0512 0.8547 1.0000
0.000 0.2833 0.01590 0.00922 -0.0489 0.8439 1.0000
0.250 0.3034 0.01602 0.00924 -0.0471 0.8311 1.0000
0.500 0.3231 0.01613 0.00925 -0.0450 0.8184 1.0000
0.750 0.3425 0.01619 0.00922 -0.0427 0.8054 1.0000
1.000 0.3616 0.01618 0.00911 -0.0401 0.7919 1.0000
1.250 0.3810 0.01612 0.00896 -0.0374 0.7791 1.0000
1.500 0.4018 0.01610 0.00888 -0.0353 0.7654 1.0000
1.750 0.4227 0.01602 0.00876 -0.0332 0.7506 1.0000
2.000 0.4435 0.01586 0.00855 -0.0309 0.7348 1.0000
2.250 0.4645 0.01562 0.00825 -0.0284 0.7176 1.0000
2.500 0.4856 0.01527 0.00785 -0.0258 0.6978 1.0000
2.750 0.5069 0.01485 0.00741 -0.0234 0.6697 1.0000
3.000 0.5280 0.01434 0.00683 -0.0207 0.6264 1.0000
3.250 0.5473 0.01399 0.00611 -0.0176 0.5203 1.0000
3.500 0.5628 0.01498 0.00598 -0.0150 0.3817 1.0000
3.750 0.5831 0.01593 0.00649 -0.0138 0.3425 1.0000
4.000 0.6050 0.01672 0.00699 -0.0128 0.3204 1.0000
4.250 0.6280 0.01745 0.00752 -0.0118 0.3046 1.0000
4.500 0.6517 0.01815 0.00808 -0.0110 0.2917 1.0000
4.750 0.6756 0.01890 0.00868 -0.0103 0.2807 1.0000
5.000 0.6997 0.01963 0.00930 -0.0096 0.2705 1.0000
5.250 0.7242 0.02038 0.01004 -0.0090 0.2621 1.0000
5.500 0.7488 0.02118 0.01073 -0.0084 0.2543 1.0000
5.750 0.7732 0.02199 0.01158 -0.0079 0.2469 1.0000
6.000 0.7976 0.02277 0.01235 -0.0073 0.2395 1.0000
6.250 0.8223 0.02379 0.01333 -0.0068 0.2337 1.0000
6.500 0.8463 0.02463 0.01432 -0.0062 0.2275 1.0000
6.750 0.8706 0.02557 0.01523 -0.0057 0.2217 1.0000
7.000 0.8939 0.02667 0.01646 -0.0051 0.2162 1.0000
7.250 0.9172 0.02767 0.01761 -0.0045 0.2107 1.0000
7.500 0.9410 0.02875 0.01870 -0.0040 0.2060 1.0000
7.750 0.9632 0.03016 0.02027 -0.0035 0.2013 1.0000
8.000 0.9843 0.03133 0.02170 -0.0027 0.1959 1.0000
8.250 1.0071 0.03250 0.02294 -0.0021 0.1915 1.0000
8.500 1.0299 0.03430 0.02470 -0.0018 0.1876 1.0000
8.750 1.0459 0.03586 0.02676 -0.0006 0.1832 1.0000
9.000 1.0645 0.03735 0.02848 0.0002 0.1784 1.0000
9.250 1.0868 0.03868 0.02981 0.0007 0.1746 1.0000
9.500 1.1042 0.04097 0.03225 0.0014 0.1712 1.0000
9.750 1.1115 0.04348 0.03532 0.0029 0.1673 1.0000
10.000 1.1231 0.04565 0.03777 0.0040 0.1632 1.0000
10.250 1.1440 0.04701 0.03913 0.0046 0.1597 1.0000
10.500 1.1608 0.04947 0.04160 0.0050 0.1565 1.0000
10.750 1.1488 0.05355 0.04633 0.0072 0.1539 1.0000
11.000 1.1334 0.05804 0.05125 0.0089 0.1516 1.0000
11.250 1.1098 0.06307 0.05659 0.0103 0.1502 1.0000
11.500 1.0535 0.07059 0.06438 0.0109 0.1513 1.0000
11.750 0.9760 0.08369 0.07762 0.0051 0.1545 1.0000
12.000 0.9246 0.09797 0.09194 -0.0028 0.1578 1.0000
12.250 0.9262 0.10301 0.09700 -0.0034 0.1569 1.0000
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Polar data table (+)
Polar graphs
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