RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 6 AIRFOIL (raf6-il) Reynolds number: 500,000 Max Cl/Cd: 90.47 at α=0.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf6-il-500000.txt Download as CSV file: xf-raf6-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3863 0.09972 0.09755 -0.0246 1.0000 0.0167
-8.500 -0.3968 0.09693 0.09480 -0.0248 1.0000 0.0169
-8.250 -0.4101 0.09450 0.09242 -0.0240 1.0000 0.0169
-8.000 -0.4095 0.09021 0.08816 -0.0289 0.9974 0.0170
-7.750 -0.3847 0.08276 0.08067 -0.0413 0.9931 0.0170
-7.500 -0.3732 0.07754 0.07545 -0.0447 0.9898 0.0173
-7.250 -0.3544 0.07396 0.07185 -0.0478 0.9874 0.0176
-7.000 -0.3311 0.06969 0.06754 -0.0534 0.9853 0.0180
-6.750 -0.3111 0.06533 0.06312 -0.0583 0.9800 0.0185
-6.500 -0.2849 0.06041 0.05812 -0.0644 0.9764 0.0193
-6.250 -0.2415 0.05275 0.05009 -0.0741 0.9736 0.0213
-6.000 -0.2245 0.04696 0.04407 -0.0761 0.9664 0.0215
-5.750 -0.2014 0.04366 0.04074 -0.0785 0.9638 0.0220
-5.500 -0.1726 0.04083 0.03783 -0.0812 0.9620 0.0225
-5.250 -0.1402 0.03782 0.03467 -0.0841 0.9607 0.0233
-5.000 -0.1185 0.03506 0.03172 -0.0839 0.9528 0.0246
-4.750 -0.0822 0.03040 0.02638 -0.0849 0.9499 0.0270
-4.500 -0.0511 0.02744 0.02339 -0.0874 0.9484 0.0279
-4.250 -0.0233 0.02573 0.02160 -0.0881 0.9435 0.0291
-4.000 0.0080 0.02392 0.01956 -0.0886 0.9386 0.0316
-2.750 0.1668 0.01267 0.00675 -0.0896 0.9088 0.0282
-2.500 0.1990 0.01202 0.00608 -0.0904 0.8987 0.0300
-2.250 0.2334 0.01119 0.00516 -0.0914 0.8904 0.0313
-2.000 0.2648 0.01064 0.00453 -0.0919 0.8802 0.0328
-1.750 0.2950 0.00993 0.00377 -0.0923 0.8690 0.0355
-1.500 0.3253 0.00959 0.00338 -0.0927 0.8566 0.0382
-1.250 0.3546 0.00933 0.00305 -0.0928 0.8439 0.0410
-1.000 0.3826 0.00901 0.00266 -0.0926 0.8319 0.0452
-0.750 0.4100 0.00886 0.00245 -0.0924 0.8196 0.0503
-0.500 0.4366 0.00871 0.00224 -0.0919 0.8067 0.0573
-0.250 0.4629 0.00857 0.00208 -0.0914 0.7926 0.0729
0.000 0.4686 0.00628 0.00223 -0.0871 0.7800 0.8801
0.250 0.5545 0.00647 0.00235 -0.0989 0.7576 0.9986
0.500 0.5819 0.00654 0.00231 -0.0988 0.7340 1.0000
0.750 0.6025 0.00666 0.00225 -0.0971 0.6995 1.0000
1.000 0.6192 0.00693 0.00222 -0.0946 0.6400 1.0000
1.250 0.6318 0.00744 0.00229 -0.0913 0.5576 1.0000
1.500 0.6458 0.00797 0.00243 -0.0885 0.4865 1.0000
1.750 0.6635 0.00837 0.00257 -0.0864 0.4411 1.0000
2.000 0.6830 0.00871 0.00270 -0.0847 0.4081 1.0000
2.250 0.7033 0.00901 0.00283 -0.0832 0.3782 1.0000
2.500 0.7244 0.00930 0.00296 -0.0819 0.3489 1.0000
2.750 0.7459 0.00959 0.00310 -0.0806 0.3220 1.0000
3.000 0.7678 0.00988 0.00325 -0.0795 0.3003 1.0000
3.250 0.7905 0.01015 0.00341 -0.0785 0.2835 1.0000
3.500 0.8133 0.01043 0.00359 -0.0775 0.2705 1.0000
3.750 0.8368 0.01067 0.00376 -0.0767 0.2603 1.0000
4.000 0.8605 0.01091 0.00396 -0.0759 0.2526 1.0000
4.250 0.8841 0.01116 0.00415 -0.0752 0.2453 1.0000
4.500 0.9079 0.01141 0.00438 -0.0744 0.2391 1.0000
4.750 0.9321 0.01163 0.00458 -0.0738 0.2335 1.0000
5.000 0.9550 0.01195 0.00484 -0.0729 0.2280 1.0000
5.250 0.9796 0.01215 0.00507 -0.0723 0.2234 1.0000
5.500 1.0038 0.01237 0.00530 -0.0717 0.2187 1.0000
5.750 1.0267 0.01269 0.00558 -0.0709 0.2142 1.0000
6.000 1.0504 0.01297 0.00586 -0.0702 0.2103 1.0000
6.250 1.0748 0.01317 0.00611 -0.0696 0.2063 1.0000
6.500 1.0982 0.01344 0.00638 -0.0689 0.2020 1.0000
6.750 1.1199 0.01387 0.00677 -0.0679 0.1974 1.0000
7.000 1.1444 0.01405 0.00703 -0.0674 0.1945 1.0000
7.250 1.1681 0.01428 0.00731 -0.0668 0.1908 1.0000
7.500 1.1909 0.01458 0.00762 -0.0660 0.1870 1.0000
7.750 1.2122 0.01501 0.00803 -0.0650 0.1825 1.0000
8.000 1.2367 0.01510 0.00821 -0.0645 0.1774 1.0000
8.250 1.2590 0.01535 0.00845 -0.0636 0.1715 1.0000
8.500 1.2808 0.01565 0.00879 -0.0627 0.1669 1.0000
8.750 1.3032 0.01584 0.00905 -0.0619 0.1618 1.0000
9.000 1.3229 0.01619 0.00937 -0.0606 0.1560 1.0000
9.250 1.3449 0.01643 0.00969 -0.0597 0.1514 1.0000
9.500 1.3655 0.01673 0.01001 -0.0586 0.1456 1.0000
9.750 1.3855 0.01708 0.01040 -0.0575 0.1398 1.0000
10.000 1.4061 0.01740 0.01076 -0.0564 0.1332 1.0000
10.250 1.4254 0.01779 0.01117 -0.0552 0.1257 1.0000
10.500 1.4438 0.01827 0.01163 -0.0539 0.1150 1.0000
10.750 1.4605 0.01888 0.01219 -0.0524 0.0987 1.0000
11.000 1.4662 0.02035 0.01334 -0.0494 0.0598 1.0000
11.250 1.4619 0.02255 0.01530 -0.0450 0.0267 1.0000
11.500 1.4692 0.02385 0.01667 -0.0422 0.0228 1.0000
11.750 1.4789 0.02493 0.01787 -0.0399 0.0207 1.0000
12.000 1.4874 0.02609 0.01914 -0.0375 0.0192 1.0000
12.250 1.4911 0.02761 0.02077 -0.0347 0.0178 1.0000
12.500 1.4913 0.02940 0.02271 -0.0318 0.0168 1.0000
12.750 1.4977 0.03077 0.02420 -0.0297 0.0162 1.0000
13.000 1.5013 0.03240 0.02596 -0.0275 0.0155 1.0000
13.250 1.5029 0.03426 0.02794 -0.0255 0.0149 1.0000
13.500 1.5020 0.03643 0.03023 -0.0236 0.0144 1.0000
13.750 1.4982 0.03899 0.03292 -0.0219 0.0140 1.0000
14.000 1.4895 0.04220 0.03625 -0.0204 0.0136 1.0000
14.250 1.4750 0.04624 0.04045 -0.0194 0.0133 1.0000
14.500 1.4570 0.05108 0.04545 -0.0190 0.0130 1.0000
14.750 1.4505 0.05494 0.04946 -0.0193 0.0128 1.0000
15.000 1.4421 0.05936 0.05403 -0.0202 0.0126 1.0000
15.250 1.4312 0.06449 0.05931 -0.0217 0.0125 1.0000
15.500 1.4178 0.07034 0.06532 -0.0239 0.0123 1.0000
15.750 1.4024 0.07674 0.07188 -0.0265 0.0122 1.0000
16.000 1.3855 0.08354 0.07884 -0.0292 0.0121 1.0000
16.250 1.3680 0.09040 0.08583 -0.0320 0.0121 1.0000
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Polar data table (+)
Polar graphs
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