RAF 15 AIRFOIL (raf15-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 15 AIRFOIL (raf15-il) Reynolds number: 500,000 Max Cl/Cd: 76.87 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf15-il-500000-n5.txt Download as CSV file: xf-raf15-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 15 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5300 0.08577 0.08352 -0.0084 1.0000 0.0072
-8.250 -0.5402 0.08220 0.07998 -0.0091 1.0000 0.0071
-8.000 -0.5501 0.07808 0.07590 -0.0100 1.0000 0.0073
-7.750 -0.5542 0.07309 0.07091 -0.0122 1.0000 0.0073
-7.500 -0.5584 0.06731 0.06511 -0.0145 1.0000 0.0073
-7.000 -0.6459 0.02332 0.01924 -0.0166 1.0000 0.0073
-6.750 -0.6341 0.02076 0.01622 -0.0140 1.0000 0.0074
-6.500 -0.6191 0.01893 0.01403 -0.0116 1.0000 0.0076
-6.250 -0.5990 0.01731 0.01209 -0.0103 0.9994 0.0080
-6.000 -0.5718 0.01630 0.01089 -0.0103 0.9980 0.0083
-5.750 -0.5428 0.01567 0.01014 -0.0107 0.9965 0.0088
-5.500 -0.5132 0.01504 0.00938 -0.0111 0.9951 0.0094
-5.250 -0.4831 0.01461 0.00882 -0.0116 0.9930 0.0103
-5.000 -0.4530 0.01382 0.00786 -0.0120 0.9907 0.0108
-4.750 -0.4206 0.01298 0.00688 -0.0129 0.9883 0.0117
-4.500 -0.3839 0.01255 0.00640 -0.0148 0.9847 0.0128
-4.250 -0.3471 0.01208 0.00587 -0.0166 0.9800 0.0143
-4.000 -0.3131 0.01168 0.00539 -0.0178 0.9758 0.0155
-3.750 -0.2833 0.01117 0.00489 -0.0180 0.9704 0.0179
-3.500 -0.2541 0.01085 0.00454 -0.0182 0.9640 0.0201
-2.750 -0.1767 0.00984 0.00344 -0.0161 0.9369 0.0274
-2.500 -0.1452 0.00955 0.00312 -0.0167 0.9298 0.0304
-2.250 -0.1134 0.00929 0.00281 -0.0173 0.9218 0.0324
-2.000 -0.0739 0.00889 0.00237 -0.0197 0.9139 0.0365
-1.750 -0.0280 0.00851 0.00199 -0.0235 0.9042 0.0421
-1.500 0.0285 0.00803 0.00166 -0.0299 0.8906 0.0844
-1.250 0.0953 0.00778 0.00147 -0.0389 0.8634 0.1299
-1.000 0.1374 0.00774 0.00140 -0.0420 0.8209 0.1603
-0.750 0.1645 0.00793 0.00134 -0.0415 0.7490 0.1770
-0.500 0.1836 0.00820 0.00133 -0.0394 0.6803 0.1875
-0.250 0.2034 0.00843 0.00133 -0.0374 0.6228 0.1977
0.000 0.2236 0.00863 0.00135 -0.0356 0.5694 0.2097
0.250 0.2449 0.00881 0.00137 -0.0341 0.5232 0.2238
0.500 0.2671 0.00897 0.00139 -0.0327 0.4818 0.2325
0.750 0.2900 0.00911 0.00141 -0.0316 0.4451 0.2415
1.000 0.3133 0.00924 0.00144 -0.0305 0.4136 0.2503
1.250 0.3368 0.00935 0.00147 -0.0295 0.3863 0.2603
1.500 0.3605 0.00944 0.00152 -0.0286 0.3652 0.2759
1.750 0.3839 0.00948 0.00158 -0.0275 0.3474 0.3061
2.000 0.4012 0.00900 0.00162 -0.0254 0.3346 0.5195
2.500 0.5718 0.00882 0.00241 -0.0518 0.2880 0.9875
2.750 0.6128 0.00903 0.00254 -0.0549 0.2680 0.9930
3.000 0.6530 0.00920 0.00263 -0.0577 0.2500 0.9971
3.250 0.6918 0.00937 0.00271 -0.0603 0.2284 1.0000
3.500 0.7144 0.00958 0.00282 -0.0591 0.2052 1.0000
3.750 0.7370 0.00980 0.00295 -0.0580 0.1857 1.0000
4.000 0.7596 0.01002 0.00312 -0.0569 0.1686 1.0000
4.250 0.7821 0.01026 0.00329 -0.0557 0.1528 1.0000
4.500 0.8047 0.01050 0.00349 -0.0546 0.1388 1.0000
4.750 0.8271 0.01076 0.00369 -0.0534 0.1243 1.0000
5.000 0.8490 0.01106 0.00394 -0.0522 0.1074 1.0000
5.250 0.8706 0.01142 0.00421 -0.0509 0.0887 1.0000
5.500 0.8920 0.01179 0.00451 -0.0496 0.0714 1.0000
5.750 0.9136 0.01215 0.00482 -0.0483 0.0596 1.0000
6.000 0.9352 0.01250 0.00517 -0.0471 0.0527 1.0000
6.250 0.9570 0.01284 0.00552 -0.0458 0.0483 1.0000
6.500 0.9791 0.01314 0.00587 -0.0446 0.0463 1.0000
6.750 1.0007 0.01348 0.00625 -0.0433 0.0445 1.0000
7.000 1.0221 0.01384 0.00666 -0.0420 0.0429 1.0000
7.250 1.0429 0.01426 0.00713 -0.0406 0.0408 1.0000
7.500 1.0633 0.01470 0.00762 -0.0392 0.0382 1.0000
7.750 1.0851 0.01499 0.00797 -0.0380 0.0366 1.0000
8.000 1.1063 0.01533 0.00837 -0.0368 0.0345 1.0000
8.250 1.1267 0.01574 0.00882 -0.0354 0.0326 1.0000
8.500 1.1463 0.01622 0.00934 -0.0339 0.0306 1.0000
8.750 1.1662 0.01667 0.00987 -0.0324 0.0288 1.0000
9.000 1.1873 0.01697 0.01026 -0.0312 0.0271 1.0000
9.250 1.2078 0.01733 0.01067 -0.0299 0.0239 1.0000
9.500 1.2270 0.01781 0.01117 -0.0284 0.0206 1.0000
9.750 1.2456 0.01833 0.01169 -0.0269 0.0158 1.0000
10.000 1.2626 0.01898 0.01237 -0.0251 0.0129 1.0000
10.250 1.2776 0.01981 0.01322 -0.0229 0.0105 1.0000
10.500 1.2935 0.02051 0.01406 -0.0209 0.0096 1.0000
10.750 1.3081 0.02130 0.01496 -0.0187 0.0088 1.0000
11.000 1.3212 0.02218 0.01594 -0.0164 0.0080 1.0000
11.250 1.3319 0.02323 0.01711 -0.0137 0.0073 1.0000
11.500 1.3442 0.02408 0.01809 -0.0113 0.0069 1.0000
11.750 1.3563 0.02490 0.01902 -0.0089 0.0064 1.0000
12.000 1.3655 0.02592 0.02016 -0.0062 0.0060 1.0000
12.250 1.3705 0.02696 0.02133 -0.0027 0.0058 1.0000
12.500 1.3721 0.02804 0.02252 0.0012 0.0056 1.0000
12.750 1.3701 0.02939 0.02400 0.0053 0.0053 1.0000
13.000 1.3680 0.03086 0.02560 0.0091 0.0052 1.0000
13.250 1.3662 0.03241 0.02730 0.0123 0.0051 1.0000
13.500 1.3648 0.03409 0.02916 0.0150 0.0049 1.0000
13.750 1.3593 0.03625 0.03148 0.0172 0.0049 1.0000
14.000 1.3521 0.03889 0.03428 0.0187 0.0048 1.0000
14.250 1.3428 0.04221 0.03777 0.0190 0.0047 1.0000
14.500 1.3321 0.04641 0.04214 0.0178 0.0047 1.0000
14.750 1.3206 0.05160 0.04751 0.0151 0.0046 1.0000
15.000 1.3005 0.05903 0.05513 0.0104 0.0047 1.0000
15.250 1.2791 0.06698 0.06326 0.0057 0.0046 1.0000
15.500 1.2525 0.07571 0.07215 0.0010 0.0048 1.0000
15.750 1.2213 0.08502 0.08159 -0.0037 0.0049 1.0000
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Polar data table (+)
Polar graphs
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