NREL's S804 Airfoil (s804-nr) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NREL's S804 Airfoil (s804-nr) Reynolds number: 50,000 Max Cl/Cd: 4.76 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-s804-nr-50000.txt Download as CSV file: xf-s804-nr-50000.csv |
XFOIL Version 6.96
Calculated polar for: NREL's S804 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2216 0.11579 0.11015 -0.0209 1.0000 0.3138
-8.250 -0.2662 0.11905 0.11367 -0.0152 1.0000 0.3175
-8.000 -0.3113 0.12211 0.11696 -0.0110 1.0000 0.3188
-7.750 -0.2538 0.11462 0.10939 -0.0168 0.9917 0.3265
-7.250 -0.4110 0.08325 0.07805 -0.0543 0.9630 0.1580
-7.000 -0.4544 0.06748 0.06166 -0.0727 0.9468 0.1422
-6.750 -0.4390 0.06030 0.05383 -0.0814 0.9326 0.1406
-6.500 -0.4117 0.05482 0.04771 -0.0880 0.9190 0.1410
-6.250 -0.3757 0.05032 0.04277 -0.0931 0.9064 0.1424
-6.000 -0.3256 0.04732 0.03967 -0.0981 0.8954 0.1493
-5.750 -0.2923 0.04450 0.03625 -0.1014 0.8809 0.1548
-5.500 -0.2577 0.04232 0.03393 -0.1034 0.8676 0.1618
-5.250 -0.2087 0.04001 0.03135 -0.1072 0.8569 0.1753
-5.000 -0.1698 0.03832 0.02958 -0.1090 0.8445 0.1917
-4.750 -0.1398 0.03712 0.02857 -0.1094 0.8316 0.2140
-4.500 -0.0853 0.03473 0.02661 -0.1131 0.8236 0.2784
-4.250 -0.0671 0.03463 0.02725 -0.1112 0.8091 0.3568
-4.000 -0.0408 0.03710 0.03023 -0.1061 0.7982 0.4394
-3.750 -0.0220 0.03931 0.03245 -0.1009 0.7860 0.4932
-3.500 -0.0173 0.04139 0.03458 -0.0938 0.7735 0.5192
-3.250 0.0092 0.04278 0.03586 -0.0889 0.7646 0.5587
-3.000 0.0064 0.04413 0.03719 -0.0831 0.7522 0.5803
-2.750 0.0276 0.04502 0.03799 -0.0777 0.7443 0.6091
-2.500 0.0216 0.04605 0.03900 -0.0726 0.7332 0.6287
-2.250 0.0391 0.04638 0.03918 -0.0695 0.7248 0.6572
-2.000 0.0523 0.04673 0.03941 -0.0665 0.7170 0.6817
-1.750 0.0314 0.04793 0.04066 -0.0604 0.7067 0.6925
-1.500 0.0652 0.04763 0.04014 -0.0598 0.7015 0.7210
-1.250 0.0314 0.04944 0.04205 -0.0539 0.6928 0.7278
-1.000 0.0320 0.05019 0.04274 -0.0514 0.6857 0.7446
-0.750 0.0595 0.04987 0.04225 -0.0497 0.6806 0.7657
-0.500 0.0509 0.05139 0.04373 -0.0481 0.6746 0.7789
-0.250 0.0381 0.05286 0.04522 -0.0451 0.6708 0.7898
0.000 0.0392 0.05394 0.04625 -0.0437 0.6677 0.8030
0.250 0.0478 0.05477 0.04700 -0.0429 0.6645 0.8169
0.500 0.0694 0.05522 0.04732 -0.0429 0.6607 0.8318
0.750 0.0909 0.05584 0.04781 -0.0428 0.6572 0.8455
1.000 0.0908 0.05728 0.04922 -0.0420 0.6569 0.8560
1.250 0.1036 0.05865 0.05053 -0.0426 0.6577 0.8677
1.500 0.0448 0.06254 0.05466 -0.0414 0.6893 0.8722
1.750 0.0702 0.06448 0.05650 -0.0436 0.6985 0.8854
2.000 -0.0165 0.06816 0.06057 -0.0421 0.8231 0.8872
2.250 -0.0097 0.06824 0.06059 -0.0414 0.8158 0.8990
2.500 0.0161 0.06938 0.06164 -0.0431 0.8065 0.9138
2.750 0.0482 0.07169 0.06385 -0.0460 0.8014 0.9301
3.000 0.0621 0.07165 0.06380 -0.0465 0.7881 0.9445
3.500 0.1183 0.07467 0.06671 -0.0534 0.7697 1.0000
3.750 0.1672 0.07835 0.07022 -0.0602 0.7632 1.0000
4.000 0.1773 0.07899 0.07080 -0.0613 0.7520 1.0000
4.250 0.2204 0.08216 0.07385 -0.0670 0.7440 1.0000
4.500 0.2450 0.08465 0.07625 -0.0704 0.7382 1.0000
4.750 0.2691 0.08627 0.07780 -0.0732 0.7261 1.0000
5.000 0.3221 0.09118 0.08257 -0.0800 0.7208 1.0000
5.250 0.3204 0.09102 0.08240 -0.0796 0.7079 1.0000
5.500 0.3638 0.09491 0.08618 -0.0848 0.7015 1.0000
5.750 0.3712 0.09631 0.08754 -0.0857 0.6921 1.0000
6.000 0.4048 0.09928 0.09043 -0.0894 0.6827 1.0000
6.250 0.4538 0.10486 0.09588 -0.0950 0.6785 1.0000
6.500 0.4430 0.10402 0.09506 -0.0934 0.6649 1.0000
6.750 0.4856 0.10854 0.09949 -0.0977 0.6593 1.0000
7.000 0.4809 0.10915 0.10008 -0.0970 0.6479 1.0000
7.250 0.5140 0.11265 0.10351 -0.0998 0.6403 1.0000
7.500 0.5279 0.11555 0.10637 -0.1011 0.6345 1.0000
7.750 0.5401 0.11715 0.10795 -0.1017 0.6225 1.0000
8.000 0.5822 0.12237 0.11310 -0.1049 0.6174 1.0000
8.250 0.5679 0.12218 0.11293 -0.1036 0.6057 1.0000
8.500 0.5991 0.12596 0.11668 -0.1056 0.5988 1.0000
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Polar data table (+)
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