RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 50,000 Max Cl/Cd: 36.66 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf6-il-50000-n5.txt Download as CSV file: xf-raf6-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3588 0.10817 0.10203 -0.0255 1.0000 0.0821
-7.500 -0.3740 0.10695 0.10094 -0.0242 1.0000 0.0833
-7.250 -0.3876 0.10559 0.09970 -0.0254 1.0000 0.0845
-7.000 -0.3980 0.10401 0.09819 -0.0279 1.0000 0.0851
-6.750 -0.3970 0.10006 0.09433 -0.0259 1.0000 0.0862
-6.500 -0.3930 0.09655 0.09087 -0.0221 1.0000 0.0885
-6.250 -0.3931 0.09387 0.08823 -0.0210 1.0000 0.0913
-6.000 -0.3936 0.09123 0.08563 -0.0217 1.0000 0.0950
-5.750 -0.3908 0.08971 0.08390 -0.0305 1.0000 0.0995
-5.500 -0.3890 0.08499 0.07940 -0.0246 1.0000 0.1023
-5.250 -0.3841 0.08211 0.07655 -0.0232 1.0000 0.1075
-4.750 -0.3625 0.07560 0.06996 -0.0269 0.9986 0.1236
-4.500 -0.3326 0.07165 0.06579 -0.0343 0.9937 0.1420
-4.250 -0.3113 0.06799 0.06212 -0.0360 0.9897 0.1564
-4.000 -0.2874 0.06469 0.05877 -0.0381 0.9856 0.1747
-3.500 -0.1786 0.05323 0.04583 -0.0527 0.9779 0.0610
-3.250 -0.1485 0.04999 0.04236 -0.0549 0.9736 0.0586
-3.000 -0.1119 0.04694 0.03892 -0.0577 0.9699 0.0568
-2.750 -0.0752 0.04412 0.03560 -0.0601 0.9658 0.0541
-2.500 -0.0360 0.04161 0.03242 -0.0621 0.9615 0.0513
-2.250 0.0028 0.03952 0.02986 -0.0644 0.9579 0.0516
-2.000 0.0353 0.03786 0.02789 -0.0656 0.9522 0.0541
-1.750 0.0771 0.03621 0.02580 -0.0680 0.9476 0.0555
-1.500 0.1136 0.03480 0.02398 -0.0692 0.9392 0.0575
-1.250 0.1533 0.03368 0.02237 -0.0707 0.9307 0.0619
-1.000 0.1950 0.03232 0.02089 -0.0727 0.9234 0.0663
-0.750 0.2287 0.03154 0.01985 -0.0733 0.9136 0.0739
-0.500 0.2671 0.03069 0.01886 -0.0750 0.9063 0.0834
-0.250 0.3016 0.02999 0.01806 -0.0761 0.8971 0.0976
0.250 0.3712 0.02633 0.01682 -0.0778 0.8797 1.0000
0.500 0.4046 0.02625 0.01628 -0.0783 0.8660 1.0000
0.750 0.4415 0.02597 0.01567 -0.0793 0.8511 1.0000
1.000 0.4791 0.02556 0.01499 -0.0802 0.8356 1.0000
1.250 0.5156 0.02511 0.01434 -0.0809 0.8201 1.0000
1.500 0.5504 0.02470 0.01376 -0.0812 0.8048 1.0000
1.750 0.5835 0.02437 0.01331 -0.0814 0.7891 1.0000
2.000 0.6147 0.02410 0.01294 -0.0813 0.7719 1.0000
2.250 0.6450 0.02383 0.01259 -0.0809 0.7530 1.0000
2.500 0.6722 0.02363 0.01233 -0.0801 0.7310 1.0000
2.750 0.6958 0.02353 0.01218 -0.0787 0.7048 1.0000
3.000 0.7214 0.02339 0.01199 -0.0776 0.6751 1.0000
3.250 0.7521 0.02312 0.01161 -0.0772 0.6416 1.0000
3.500 0.7856 0.02288 0.01117 -0.0772 0.6022 1.0000
3.750 0.8181 0.02285 0.01083 -0.0772 0.5592 1.0000
4.000 0.8453 0.02316 0.01081 -0.0765 0.5180 1.0000
4.250 0.8692 0.02371 0.01105 -0.0756 0.4830 1.0000
4.500 0.8918 0.02437 0.01146 -0.0746 0.4545 1.0000
4.750 0.9141 0.02509 0.01197 -0.0737 0.4306 1.0000
5.000 0.9370 0.02582 0.01256 -0.0730 0.4104 1.0000
5.250 0.9602 0.02655 0.01319 -0.0724 0.3924 1.0000
5.500 0.9845 0.02728 0.01384 -0.0720 0.3770 1.0000
5.750 1.0091 0.02800 0.01455 -0.0716 0.3631 1.0000
6.000 1.0336 0.02872 0.01531 -0.0713 0.3501 1.0000
6.250 1.0591 0.02947 0.01610 -0.0712 0.3389 1.0000
6.750 1.1092 0.03100 0.01774 -0.0708 0.3189 1.0000
7.000 1.1347 0.03179 0.01857 -0.0707 0.3104 1.0000
7.250 1.1585 0.03264 0.01957 -0.0703 0.3018 1.0000
7.500 1.1842 0.03351 0.02049 -0.0702 0.2946 1.0000
7.750 1.2061 0.03445 0.02164 -0.0696 0.2866 1.0000
8.000 1.2313 0.03536 0.02258 -0.0695 0.2799 1.0000
8.250 1.2508 0.03643 0.02393 -0.0686 0.2726 1.0000
8.500 1.2754 0.03741 0.02502 -0.0684 0.2668 1.0000
8.750 1.2940 0.03863 0.02652 -0.0674 0.2605 1.0000
9.000 1.3139 0.03977 0.02786 -0.0665 0.2545 1.0000
9.250 1.3371 0.04092 0.02913 -0.0662 0.2494 1.0000
9.500 1.3503 0.04244 0.03108 -0.0645 0.2441 1.0000
9.750 1.3689 0.04376 0.03263 -0.0635 0.2392 1.0000
10.000 1.3918 0.04504 0.03405 -0.0632 0.2351 1.0000
10.250 1.3968 0.04694 0.03642 -0.0605 0.2305 1.0000
10.500 1.4078 0.04867 0.03849 -0.0587 0.2265 1.0000
10.750 1.4246 0.05026 0.04034 -0.0577 0.2233 1.0000
11.000 1.4340 0.05175 0.04213 -0.0556 0.2186 1.0000
11.250 1.4237 0.05352 0.04426 -0.0511 0.2127 1.0000
11.500 1.4256 0.05329 0.04403 -0.0476 0.2027 1.0000
11.750 1.4112 0.05483 0.04584 -0.0430 0.1955 1.0000
12.000 1.4056 0.05532 0.04639 -0.0395 0.1865 1.0000
12.250 1.3906 0.05757 0.04894 -0.0362 0.1801 1.0000
12.500 1.3839 0.05878 0.05024 -0.0336 0.1722 1.0000
12.750 1.3668 0.06232 0.05414 -0.0316 0.1675 1.0000
13.000 1.3614 0.06419 0.05615 -0.0302 0.1608 1.0000
13.250 1.3418 0.06882 0.06109 -0.0295 0.1571 1.0000
13.500 1.3174 0.07466 0.06723 -0.0298 0.1547 1.0000
13.750 1.2872 0.08214 0.07493 -0.0312 0.1542 1.0000
14.000 1.2448 0.09287 0.08580 -0.0347 0.1568 1.0000
14.250 1.2010 0.10559 0.09856 -0.0397 0.1606 1.0000
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Polar data table (+)
Polar graphs
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