RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 1,000,000 Max Cl/Cd: 104.97 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf6-il-1000000-n5.txt Download as CSV file: xf-raf6-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 6 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.1917 0.08979 0.08809 -0.0713 0.9847 0.0060
-8.750 -0.1829 0.08612 0.08442 -0.0740 0.9824 0.0060
-8.500 -0.1720 0.08222 0.08053 -0.0775 0.9807 0.0060
-8.000 -0.1455 0.07543 0.07375 -0.0846 0.9782 0.0056
-7.750 -0.1415 0.07203 0.07037 -0.0871 0.9721 0.0054
-7.500 -0.1252 0.06714 0.06547 -0.0942 0.9691 0.0052
-7.250 -0.1144 0.06280 0.06112 -0.0989 0.9626 0.0051
-7.000 -0.0948 0.05782 0.05609 -0.1054 0.9586 0.0051
-6.750 -0.0807 0.05342 0.05163 -0.1091 0.9501 0.0050
-6.500 -0.0621 0.04879 0.04690 -0.1131 0.9417 0.0050
-6.250 -0.0402 0.04447 0.04247 -0.1165 0.9323 0.0052
-6.000 -0.0173 0.03996 0.03779 -0.1194 0.9189 0.0054
-5.500 0.0066 0.01311 0.00875 -0.1190 0.8755 0.0059
-5.250 0.0318 0.01166 0.00688 -0.1185 0.8580 0.0060
-5.000 0.0570 0.01092 0.00592 -0.1180 0.8413 0.0064
-4.750 0.0831 0.01069 0.00556 -0.1176 0.8270 0.0066
-4.500 0.1091 0.01036 0.00510 -0.1172 0.8157 0.0069
-4.250 0.1353 0.00991 0.00452 -0.1168 0.8071 0.0073
-3.750 0.1875 0.00919 0.00355 -0.1160 0.7879 0.0080
-3.500 0.2133 0.00887 0.00312 -0.1155 0.7763 0.0083
-3.250 0.2392 0.00868 0.00286 -0.1150 0.7631 0.0088
-3.000 0.2652 0.00855 0.00266 -0.1146 0.7492 0.0095
-2.750 0.2919 0.00842 0.00246 -0.1143 0.7372 0.0103
-2.500 0.3182 0.00824 0.00222 -0.1139 0.7240 0.0111
-2.250 0.3439 0.00816 0.00206 -0.1134 0.7054 0.0120
-2.000 0.3690 0.00813 0.00193 -0.1127 0.6812 0.0131
-1.750 0.3932 0.00816 0.00181 -0.1119 0.6503 0.0146
-1.500 0.4165 0.00828 0.00177 -0.1109 0.6090 0.0164
-1.250 0.4390 0.00850 0.00176 -0.1098 0.5582 0.0180
-1.000 0.4612 0.00875 0.00177 -0.1087 0.5002 0.0211
-0.750 0.4844 0.00901 0.00181 -0.1078 0.4527 0.0241
-0.500 0.5091 0.00916 0.00181 -0.1071 0.4221 0.0288
-0.250 0.5347 0.00925 0.00183 -0.1067 0.4011 0.0358
0.000 0.5604 0.00933 0.00185 -0.1063 0.3825 0.0471
0.250 0.5862 0.00941 0.00188 -0.1059 0.3644 0.0633
0.500 0.6089 0.00851 0.00202 -0.1057 0.3457 0.5146
0.750 0.6266 0.00777 0.00219 -0.1037 0.3261 0.8636
1.000 0.6607 0.00784 0.00233 -0.1049 0.3007 0.9928
1.250 0.6917 0.00806 0.00240 -0.1058 0.2816 1.0000
1.500 0.7158 0.00824 0.00248 -0.1050 0.2674 1.0000
1.750 0.7404 0.00841 0.00256 -0.1044 0.2562 1.0000
2.000 0.7651 0.00857 0.00264 -0.1038 0.2466 1.0000
2.250 0.7898 0.00875 0.00274 -0.1033 0.2370 1.0000
2.500 0.8151 0.00890 0.00284 -0.1028 0.2300 1.0000
2.750 0.8403 0.00906 0.00294 -0.1023 0.2240 1.0000
3.000 0.8657 0.00921 0.00305 -0.1019 0.2194 1.0000
3.250 0.8912 0.00935 0.00316 -0.1015 0.2153 1.0000
3.500 0.9164 0.00951 0.00328 -0.1011 0.2103 1.0000
3.750 0.9414 0.00969 0.00342 -0.1006 0.2053 1.0000
4.000 0.9671 0.00982 0.00355 -0.1003 0.2026 1.0000
4.250 0.9924 0.00997 0.00369 -0.0999 0.1988 1.0000
4.500 1.0174 0.01015 0.00383 -0.0994 0.1949 1.0000
4.750 1.0421 0.01034 0.00400 -0.0990 0.1905 1.0000
5.000 1.0672 0.01049 0.00416 -0.0986 0.1878 1.0000
5.250 1.0923 0.01064 0.00432 -0.0981 0.1852 1.0000
5.500 1.1170 0.01082 0.00449 -0.0977 0.1817 1.0000
5.750 1.1412 0.01102 0.00468 -0.0971 0.1777 1.0000
6.000 1.1653 0.01123 0.00488 -0.0966 0.1740 1.0000
6.250 1.1898 0.01140 0.00507 -0.0961 0.1720 1.0000
6.500 1.2140 0.01158 0.00527 -0.0955 0.1693 1.0000
6.750 1.2376 0.01179 0.00549 -0.0949 0.1658 1.0000
7.000 1.2605 0.01204 0.00573 -0.0942 0.1610 1.0000
7.250 1.2831 0.01229 0.00596 -0.0934 0.1531 1.0000
7.500 1.3040 0.01260 0.00623 -0.0923 0.1439 1.0000
7.750 1.3241 0.01296 0.00653 -0.0910 0.1332 1.0000
8.000 1.3444 0.01331 0.00686 -0.0898 0.1235 1.0000
8.250 1.3629 0.01378 0.00726 -0.0883 0.1096 1.0000
8.500 1.3798 0.01437 0.00774 -0.0866 0.0937 1.0000
8.750 1.3949 0.01508 0.00834 -0.0846 0.0760 1.0000
9.000 1.3964 0.01680 0.00975 -0.0804 0.0221 1.0000
9.250 1.4109 0.01757 0.01050 -0.0784 0.0129 1.0000
9.500 1.4277 0.01815 0.01110 -0.0767 0.0096 1.0000
9.750 1.4446 0.01870 0.01168 -0.0752 0.0078 1.0000
10.000 1.4610 0.01929 0.01230 -0.0735 0.0066 1.0000
10.250 1.4774 0.01986 0.01293 -0.0720 0.0058 1.0000
10.500 1.4925 0.02052 0.01363 -0.0702 0.0051 1.0000
10.750 1.5081 0.02115 0.01431 -0.0686 0.0046 1.0000
11.000 1.5226 0.02184 0.01505 -0.0669 0.0042 1.0000
11.250 1.5356 0.02263 0.01590 -0.0650 0.0038 1.0000
11.500 1.5492 0.02338 0.01672 -0.0632 0.0035 1.0000
12.000 1.5734 0.02510 0.01859 -0.0595 0.0031 1.0000
12.250 1.5836 0.02611 0.01967 -0.0575 0.0029 1.0000
12.500 1.5918 0.02730 0.02093 -0.0554 0.0027 1.0000
12.750 1.6010 0.02842 0.02214 -0.0535 0.0026 1.0000
13.000 1.6091 0.02964 0.02346 -0.0517 0.0025 1.0000
13.250 1.6164 0.03098 0.02488 -0.0498 0.0023 1.0000
13.500 1.6228 0.03241 0.02640 -0.0481 0.0022 1.0000
13.750 1.6282 0.03397 0.02805 -0.0464 0.0021 1.0000
14.000 1.6321 0.03573 0.02989 -0.0448 0.0020 1.0000
14.250 1.6340 0.03772 0.03197 -0.0432 0.0019 1.0000
14.500 1.6322 0.04012 0.03449 -0.0415 0.0018 1.0000
14.750 1.6301 0.04268 0.03716 -0.0402 0.0018 1.0000
15.000 1.6283 0.04535 0.03994 -0.0391 0.0017 1.0000
15.250 1.6248 0.04831 0.04305 -0.0383 0.0017 1.0000
15.500 1.6192 0.05169 0.04655 -0.0377 0.0017 1.0000
15.750 1.6115 0.05555 0.05054 -0.0376 0.0016 1.0000
16.000 1.6018 0.05991 0.05504 -0.0379 0.0016 1.0000
16.250 1.5886 0.06502 0.06029 -0.0387 0.0016 1.0000
16.500 1.5729 0.07086 0.06629 -0.0402 0.0016 1.0000
16.750 1.5538 0.07762 0.07321 -0.0424 0.0016 1.0000
17.000 1.5308 0.08529 0.08105 -0.0452 0.0016 1.0000
17.250 1.5029 0.09395 0.08988 -0.0484 0.0016 1.0000
17.500 1.4737 0.10290 0.09899 -0.0519 0.0016 1.0000
17.750 1.4421 0.11230 0.10855 -0.0556 0.0016 1.0000
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