RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 500,000 Max Cl/Cd: 84.99 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf6-il-500000-n5.txt Download as CSV file: xf-raf6-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.2107 0.08767 0.08541 -0.0624 0.9812 0.0094
-8.000 -0.1978 0.08372 0.08146 -0.0666 0.9790 0.0098
-7.750 -0.1902 0.07994 0.07771 -0.0700 0.9740 0.0102
-7.500 -0.1732 0.07455 0.07231 -0.0786 0.9695 0.0105
-7.250 -0.1590 0.06967 0.06741 -0.0851 0.9628 0.0105
-6.750 -0.1182 0.05941 0.05704 -0.0976 0.9526 0.0106
-6.250 -0.0485 0.03521 0.03293 -0.1004 0.9264 0.0112
-5.750 -0.0125 0.02636 0.02383 -0.1064 0.9110 0.0099
-5.250 0.0055 0.03308 0.02984 -0.1118 0.9135 0.0079
-5.000 0.0346 0.02856 0.02498 -0.1133 0.9070 0.0080
-4.750 0.0613 0.02343 0.01939 -0.1135 0.8974 0.0081
-4.500 0.0923 0.02176 0.01751 -0.1149 0.8889 0.0085
-4.250 0.1253 0.01992 0.01539 -0.1163 0.8786 0.0091
-4.000 0.1543 0.01578 0.01063 -0.1160 0.8666 0.0094
-3.750 0.1836 0.01291 0.00716 -0.1158 0.8541 0.0101
-3.500 0.2125 0.01179 0.00569 -0.1157 0.8417 0.0106
-3.000 0.2677 0.01110 0.00479 -0.1156 0.8177 0.0124
-2.750 0.2938 0.01062 0.00415 -0.1150 0.8040 0.0132
-2.500 0.3195 0.01022 0.00361 -0.1144 0.7897 0.0141
-2.250 0.3456 0.01000 0.00334 -0.1140 0.7779 0.0153
-2.000 0.3722 0.00984 0.00314 -0.1137 0.7689 0.0171
-1.750 0.3985 0.00958 0.00280 -0.1133 0.7597 0.0187
-1.500 0.4247 0.00942 0.00261 -0.1129 0.7481 0.0211
-1.250 0.4506 0.00934 0.00246 -0.1124 0.7335 0.0241
-1.000 0.4762 0.00918 0.00222 -0.1118 0.7166 0.0284
-0.750 0.5014 0.00914 0.00210 -0.1112 0.6948 0.0343
-0.500 0.5251 0.00917 0.00200 -0.1102 0.6642 0.0415
-0.250 0.5464 0.00933 0.00196 -0.1088 0.6193 0.0530
0.000 0.5655 0.00960 0.00199 -0.1069 0.5593 0.0803
0.250 0.5751 0.00809 0.00228 -0.1041 0.5061 0.7932
0.500 0.6161 0.00829 0.00250 -0.1066 0.4512 1.0000
0.750 0.6376 0.00856 0.00257 -0.1054 0.4240 1.0000
1.000 0.6601 0.00880 0.00264 -0.1043 0.4016 1.0000
1.250 0.6830 0.00903 0.00272 -0.1033 0.3808 1.0000
1.500 0.7062 0.00927 0.00281 -0.1024 0.3595 1.0000
1.750 0.7295 0.00951 0.00291 -0.1016 0.3371 1.0000
2.000 0.7528 0.00976 0.00302 -0.1007 0.3160 1.0000
2.250 0.7764 0.01001 0.00314 -0.1000 0.2981 1.0000
2.500 0.8001 0.01024 0.00327 -0.0992 0.2827 1.0000
2.750 0.8241 0.01047 0.00340 -0.0986 0.2699 1.0000
3.000 0.8481 0.01070 0.00355 -0.0979 0.2587 1.0000
3.250 0.8719 0.01094 0.00371 -0.0972 0.2488 1.0000
3.500 0.8962 0.01115 0.00387 -0.0966 0.2410 1.0000
3.750 0.9202 0.01138 0.00405 -0.0960 0.2349 1.0000
4.000 0.9447 0.01157 0.00422 -0.0954 0.2297 1.0000
4.250 0.9686 0.01181 0.00441 -0.0947 0.2239 1.0000
4.500 0.9924 0.01203 0.00462 -0.0941 0.2190 1.0000
4.750 1.0167 0.01223 0.00482 -0.0935 0.2148 1.0000
5.000 1.0403 0.01247 0.00504 -0.0929 0.2102 1.0000
5.250 1.0633 0.01274 0.00528 -0.0921 0.2057 1.0000
5.500 1.0874 0.01293 0.00550 -0.0915 0.2027 1.0000
5.750 1.1110 0.01315 0.00574 -0.0908 0.1988 1.0000
6.000 1.1340 0.01339 0.00599 -0.0901 0.1948 1.0000
6.250 1.1563 0.01367 0.00626 -0.0892 0.1912 1.0000
6.500 1.1785 0.01391 0.00654 -0.0883 0.1881 1.0000
6.750 1.2009 0.01413 0.00680 -0.0874 0.1849 1.0000
7.000 1.2228 0.01439 0.00708 -0.0865 0.1812 1.0000
7.250 1.2442 0.01467 0.00740 -0.0854 0.1778 1.0000
7.500 1.2648 0.01501 0.00775 -0.0843 0.1744 1.0000
7.750 1.2869 0.01525 0.00806 -0.0834 0.1707 1.0000
8.000 1.3077 0.01558 0.00837 -0.0824 0.1623 1.0000
8.250 1.3284 0.01591 0.00871 -0.0813 0.1533 1.0000
8.500 1.3480 0.01632 0.00909 -0.0802 0.1442 1.0000
8.750 1.3667 0.01679 0.00951 -0.0788 0.1319 1.0000
9.000 1.3843 0.01734 0.01002 -0.0774 0.1181 1.0000
9.250 1.4001 0.01801 0.01062 -0.0757 0.1012 1.0000
9.500 1.4119 0.01899 0.01144 -0.0734 0.0786 1.0000
9.750 1.4075 0.02121 0.01329 -0.0688 0.0210 1.0000
10.000 1.4187 0.02222 0.01434 -0.0665 0.0136 1.0000
10.250 1.4319 0.02307 0.01525 -0.0646 0.0109 1.0000
10.500 1.4455 0.02389 0.01616 -0.0627 0.0093 1.0000
10.750 1.4581 0.02478 0.01714 -0.0609 0.0082 1.0000
11.000 1.4702 0.02570 0.01815 -0.0590 0.0072 1.0000
11.250 1.4806 0.02677 0.01931 -0.0569 0.0066 1.0000
11.500 1.4910 0.02783 0.02048 -0.0550 0.0061 1.0000
11.750 1.4999 0.02902 0.02178 -0.0530 0.0057 1.0000
12.000 1.5070 0.03037 0.02323 -0.0509 0.0053 1.0000
12.500 1.5181 0.03343 0.02654 -0.0469 0.0047 1.0000
12.750 1.5227 0.03509 0.02832 -0.0451 0.0044 1.0000
13.000 1.5262 0.03691 0.03025 -0.0434 0.0042 1.0000
13.250 1.5274 0.03901 0.03247 -0.0418 0.0040 1.0000
13.500 1.5248 0.04154 0.03512 -0.0401 0.0039 1.0000
13.750 1.5204 0.04438 0.03811 -0.0387 0.0038 1.0000
14.000 1.5165 0.04733 0.04120 -0.0377 0.0037 1.0000
14.250 1.5103 0.05070 0.04473 -0.0369 0.0036 1.0000
14.500 1.5021 0.05455 0.04874 -0.0366 0.0035 1.0000
14.750 1.4919 0.05895 0.05329 -0.0368 0.0034 1.0000
15.000 1.4795 0.06396 0.05846 -0.0375 0.0034 1.0000
15.250 1.4651 0.06964 0.06431 -0.0390 0.0033 1.0000
15.500 1.4479 0.07619 0.07103 -0.0411 0.0033 1.0000
15.750 1.4288 0.08338 0.07840 -0.0438 0.0033 1.0000
16.000 1.4073 0.09112 0.08631 -0.0468 0.0033 1.0000
16.250 1.3843 0.09918 0.09453 -0.0499 0.0033 1.0000
16.500 1.3616 0.10717 0.10267 -0.0530 0.0033 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAF 6 AIRFOIL (raf6-il)