NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Reynolds number: 50,000 Max Cl/Cd: 31.23 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08b3-il-50000-n5.txt Download as CSV file: xf-rc08b3-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(B)3 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5877 0.09879 0.09169 -0.0116 1.0000 0.0682
-8.750 -0.5855 0.09439 0.08732 -0.0132 1.0000 0.0667
-8.250 -0.6107 0.08226 0.07503 -0.0233 1.0000 0.0566
-8.000 -0.6086 0.07800 0.07070 -0.0238 1.0000 0.0564
-7.750 -0.6058 0.07380 0.06642 -0.0240 1.0000 0.0561
-7.500 -0.6017 0.06969 0.06220 -0.0240 1.0000 0.0558
-7.250 -0.5962 0.06569 0.05808 -0.0238 1.0000 0.0552
-7.000 -0.5898 0.06176 0.05400 -0.0233 1.0000 0.0544
-6.750 -0.5827 0.05790 0.04993 -0.0225 1.0000 0.0536
-6.500 -0.5744 0.05414 0.04591 -0.0215 1.0000 0.0528
-6.250 -0.5647 0.05053 0.04199 -0.0203 1.0000 0.0522
-6.000 -0.5536 0.04711 0.03821 -0.0187 1.0000 0.0518
-5.750 -0.5410 0.04391 0.03462 -0.0170 1.0000 0.0515
-5.500 -0.5268 0.04097 0.03123 -0.0152 1.0000 0.0514
-5.250 -0.5110 0.03826 0.02810 -0.0133 1.0000 0.0515
-5.000 -0.4937 0.03583 0.02525 -0.0115 1.0000 0.0525
-4.750 -0.4747 0.03374 0.02268 -0.0098 1.0000 0.0549
-4.500 -0.4543 0.03183 0.02031 -0.0082 1.0000 0.0570
-4.250 -0.4332 0.02987 0.01820 -0.0070 1.0000 0.0589
-4.000 -0.4104 0.02820 0.01634 -0.0059 1.0000 0.0606
-3.750 -0.3864 0.02671 0.01461 -0.0049 1.0000 0.0630
-3.500 -0.3610 0.02536 0.01307 -0.0040 1.0000 0.0663
-3.250 -0.3351 0.02419 0.01173 -0.0033 1.0000 0.0717
-3.000 -0.3105 0.02329 0.01073 -0.0026 1.0000 0.0822
-2.750 -0.2868 0.02228 0.00970 -0.0018 1.0000 0.0941
-2.250 -0.0944 0.01739 0.00770 -0.0255 1.0000 1.0000
-2.000 -0.0796 0.01732 0.00740 -0.0234 1.0000 1.0000
-1.750 -0.0657 0.01728 0.00717 -0.0212 1.0000 1.0000
-1.500 -0.0524 0.01728 0.00699 -0.0189 1.0000 1.0000
-1.250 -0.0396 0.01731 0.00687 -0.0165 1.0000 1.0000
-1.000 -0.0271 0.01737 0.00677 -0.0140 1.0000 1.0000
-0.750 -0.0148 0.01746 0.00673 -0.0115 1.0000 1.0000
-0.500 -0.0025 0.01757 0.00674 -0.0091 1.0000 1.0000
-0.250 0.0099 0.01771 0.00678 -0.0067 1.0000 1.0000
0.000 0.0225 0.01788 0.00686 -0.0044 1.0000 1.0000
0.250 0.0353 0.01807 0.00696 -0.0022 1.0000 1.0000
0.500 0.0484 0.01829 0.00712 0.0000 1.0000 1.0000
0.750 0.0618 0.01854 0.00732 0.0019 1.0000 1.0000
1.000 0.1035 0.01885 0.00760 -0.0017 0.9895 1.0000
1.250 0.1465 0.01916 0.00792 -0.0055 0.9781 1.0000
1.500 0.1898 0.01945 0.00825 -0.0092 0.9657 1.0000
1.750 0.2343 0.01970 0.00860 -0.0130 0.9523 1.0000
2.000 0.2802 0.01990 0.00892 -0.0170 0.9380 1.0000
2.250 0.3268 0.02004 0.00921 -0.0209 0.9226 1.0000
2.500 0.3711 0.02009 0.00947 -0.0240 0.9053 1.0000
2.750 0.4119 0.02007 0.00965 -0.0261 0.8859 1.0000
3.000 0.4462 0.02002 0.00980 -0.0268 0.8631 1.0000
3.250 0.4783 0.01991 0.00993 -0.0267 0.8379 1.0000
3.500 0.5075 0.01973 0.00995 -0.0258 0.8084 1.0000
3.750 0.5348 0.01949 0.00988 -0.0241 0.7736 1.0000
4.000 0.5582 0.01928 0.00979 -0.0217 0.7283 1.0000
4.250 0.5794 0.01912 0.00968 -0.0186 0.6647 1.0000
4.500 0.5987 0.01917 0.00945 -0.0151 0.5702 1.0000
4.750 0.6137 0.01981 0.00948 -0.0115 0.4487 1.0000
5.000 0.6262 0.02101 0.01000 -0.0086 0.3284 1.0000
5.250 0.6394 0.02255 0.01089 -0.0065 0.2218 1.0000
5.500 0.6543 0.02418 0.01203 -0.0048 0.1573 1.0000
5.750 0.6718 0.02564 0.01335 -0.0033 0.1250 1.0000
6.000 0.6906 0.02697 0.01464 -0.0019 0.1054 1.0000
6.250 0.7109 0.02835 0.01608 -0.0005 0.0940 1.0000
6.500 0.7315 0.02972 0.01750 0.0007 0.0838 1.0000
6.750 0.7529 0.03119 0.01902 0.0019 0.0767 1.0000
7.000 0.7769 0.03305 0.02106 0.0029 0.0725 1.0000
7.250 0.7993 0.03487 0.02322 0.0040 0.0673 1.0000
7.500 0.8189 0.03675 0.02509 0.0048 0.0625 1.0000
7.750 0.8387 0.03913 0.02796 0.0062 0.0596 1.0000
8.000 0.8563 0.04191 0.03122 0.0077 0.0576 1.0000
8.250 0.8707 0.04494 0.03471 0.0094 0.0561 1.0000
8.500 0.8817 0.04810 0.03830 0.0112 0.0547 1.0000
8.750 0.8920 0.05081 0.04127 0.0127 0.0525 1.0000
9.000 0.9015 0.05370 0.04428 0.0139 0.0502 1.0000
9.250 0.8994 0.05742 0.04852 0.0161 0.0490 1.0000
9.500 0.8938 0.06143 0.05297 0.0182 0.0484 1.0000
9.750 0.8850 0.06557 0.05745 0.0199 0.0483 1.0000
10.000 0.8720 0.06976 0.06190 0.0213 0.0483 1.0000
10.250 0.8545 0.07384 0.06615 0.0226 0.0484 1.0000
10.500 0.8352 0.07833 0.07076 0.0225 0.0485 1.0000
10.750 0.8162 0.08361 0.07613 0.0206 0.0488 1.0000
11.000 0.7986 0.08973 0.08230 0.0173 0.0491 1.0000
11.250 0.7840 0.09649 0.08908 0.0131 0.0494 1.0000
11.500 0.7734 0.10344 0.09598 0.0090 0.0497 1.0000
11.750 0.7664 0.11011 0.10263 0.0055 0.0500 1.0000
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Polar data table (+)
Polar graphs
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