NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 50,000 Max Cl/Cd: 33.24 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08n1-il-50000.txt Download as CSV file: xf-rc08n1-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(N)1 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5845 0.10140 0.09542 0.0184 1.0000 0.2269
-7.750 -0.5878 0.09816 0.09227 0.0170 1.0000 0.2424
-7.500 -0.5956 0.09509 0.08930 0.0145 1.0000 0.2575
-7.250 -0.5900 0.09163 0.08590 0.0161 1.0000 0.2855
-7.000 -0.5751 0.08821 0.08247 0.0214 1.0000 0.3266
-6.750 -0.5512 0.08468 0.07894 0.0278 1.0000 0.3795
-6.500 -0.5394 0.08208 0.07638 0.0331 1.0000 0.4376
-6.250 -0.4984 0.07837 0.07262 0.0412 1.0000 0.5297
-6.000 -0.4583 0.07455 0.06874 0.0472 1.0000 0.6305
-5.750 -0.4129 0.07023 0.06433 0.0512 1.0000 0.7469
-4.500 -0.4570 0.05464 0.04936 0.0398 1.0000 0.6106
-4.250 -0.4191 0.04032 0.03295 -0.0150 1.0000 0.2459
-4.000 -0.3833 0.03712 0.02885 -0.0161 1.0000 0.1816
-3.750 -0.3580 0.03486 0.02593 -0.0149 1.0000 0.1562
-3.500 -0.3369 0.03301 0.02354 -0.0131 1.0000 0.1439
-3.250 -0.3165 0.03094 0.02106 -0.0115 1.0000 0.1352
-3.000 -0.2947 0.02930 0.01906 -0.0099 1.0000 0.1271
-2.750 -0.2719 0.02789 0.01717 -0.0082 1.0000 0.1197
-2.500 -0.2487 0.02658 0.01557 -0.0069 1.0000 0.1164
-2.250 -0.2239 0.02514 0.01400 -0.0060 1.0000 0.1146
-2.000 -0.1977 0.02396 0.01268 -0.0052 1.0000 0.1143
-1.750 -0.1715 0.02314 0.01159 -0.0045 1.0000 0.1179
-1.500 -0.1487 0.02213 0.01060 -0.0038 1.0000 0.1266
-1.250 -0.1273 0.02145 0.00981 -0.0028 1.0000 0.1342
-1.000 -0.1049 0.02081 0.00911 -0.0022 1.0000 0.1476
-0.750 -0.0357 0.01720 0.00825 -0.0079 1.0000 1.0000
-0.500 -0.0215 0.01748 0.00799 -0.0060 1.0000 1.0000
-0.250 -0.0067 0.01777 0.00794 -0.0045 1.0000 1.0000
0.000 0.0091 0.01810 0.00800 -0.0034 1.0000 1.0000
0.250 0.0255 0.01846 0.00813 -0.0025 1.0000 1.0000
0.500 0.0424 0.01885 0.00833 -0.0017 1.0000 1.0000
0.750 0.0598 0.01929 0.00860 -0.0011 1.0000 1.0000
1.000 0.0774 0.01976 0.00891 -0.0006 1.0000 1.0000
1.250 0.0952 0.02027 0.00931 -0.0003 1.0000 1.0000
1.500 0.1131 0.02083 0.00978 0.0000 1.0000 1.0000
1.750 0.1530 0.02164 0.01053 -0.0041 0.9911 1.0000
2.000 0.2015 0.02258 0.01145 -0.0097 0.9777 1.0000
2.250 0.2488 0.02352 0.01241 -0.0150 0.9633 1.0000
2.500 0.2951 0.02444 0.01339 -0.0199 0.9478 1.0000
2.750 0.3416 0.02536 0.01445 -0.0246 0.9314 1.0000
3.000 0.3923 0.02626 0.01552 -0.0298 0.9142 1.0000
3.250 0.4311 0.02712 0.01654 -0.0327 0.8950 1.0000
3.500 0.4793 0.02791 0.01757 -0.0369 0.8746 1.0000
3.750 0.5200 0.02864 0.01862 -0.0392 0.8521 1.0000
4.000 0.5632 0.02908 0.01936 -0.0408 0.8263 1.0000
4.250 0.6004 0.02911 0.01969 -0.0400 0.7951 1.0000
4.500 0.6346 0.02862 0.01955 -0.0372 0.7630 1.0000
4.750 0.6595 0.02807 0.01925 -0.0331 0.7311 1.0000
5.000 0.6807 0.02728 0.01870 -0.0282 0.6963 1.0000
5.250 0.7013 0.02595 0.01759 -0.0224 0.6578 1.0000
5.500 0.7210 0.02403 0.01584 -0.0157 0.6100 1.0000
5.750 0.7367 0.02216 0.01399 -0.0085 0.5260 1.0000
6.000 0.7448 0.02269 0.01331 -0.0021 0.3636 1.0000
6.250 0.7577 0.02517 0.01487 0.0006 0.2672 1.0000
6.500 0.7769 0.02747 0.01678 0.0024 0.2140 1.0000
6.750 0.7990 0.02983 0.01884 0.0038 0.1785 1.0000
7.000 0.8236 0.03246 0.02152 0.0049 0.1554 1.0000
7.250 0.8462 0.03532 0.02450 0.0059 0.1375 1.0000
7.500 0.8686 0.03858 0.02799 0.0069 0.1253 1.0000
7.750 0.8872 0.04224 0.03245 0.0082 0.1197 1.0000
8.000 0.9044 0.04580 0.03619 0.0091 0.1107 1.0000
8.250 0.9150 0.05000 0.04111 0.0103 0.1077 1.0000
8.500 0.9223 0.05487 0.04656 0.0112 0.1078 1.0000
8.750 0.9256 0.06006 0.05221 0.0118 0.1092 1.0000
9.000 0.9274 0.06543 0.05790 0.0122 0.1108 1.0000
9.250 0.9285 0.07090 0.06359 0.0123 0.1118 1.0000
9.500 0.8869 0.07756 0.07080 0.0095 0.1206 1.0000
9.750 0.8829 0.08341 0.07670 0.0086 0.1245 1.0000
10.000 0.8304 0.09200 0.08530 -0.0002 0.1314 1.0000
10.250 0.8048 0.10278 0.09597 -0.0094 0.1446 1.0000
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Polar data table (+)
Polar graphs
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