RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAF 28 AIRFOIL (raf28-il) Reynolds number: 100,000 Max Cl/Cd: 50.86 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-raf28-il-100000-n5.txt Download as CSV file: xf-raf28-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAF 28 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5128 0.09145 0.08634 -0.0306 1.0000 0.0324
-9.750 -0.5230 0.08502 0.07998 -0.0340 1.0000 0.0320
-9.500 -0.5619 0.06887 0.06390 -0.0471 1.0000 0.0302
-9.250 -0.5843 0.06354 0.05845 -0.0501 1.0000 0.0298
-9.000 -0.6026 0.05924 0.05404 -0.0497 1.0000 0.0297
-8.750 -0.6164 0.05499 0.04960 -0.0490 1.0000 0.0297
-8.250 -0.6343 0.04641 0.04048 -0.0458 1.0000 0.0302
-8.000 -0.6341 0.04292 0.03674 -0.0439 1.0000 0.0308
-7.750 -0.6266 0.04078 0.03447 -0.0423 1.0000 0.0322
-7.500 -0.6172 0.03885 0.03237 -0.0406 1.0000 0.0339
-7.250 -0.6088 0.03640 0.02960 -0.0385 1.0000 0.0357
-7.000 -0.6001 0.03357 0.02632 -0.0362 1.0000 0.0369
-6.750 -0.5888 0.03111 0.02339 -0.0339 1.0000 0.0384
-6.500 -0.5751 0.02944 0.02120 -0.0317 1.0000 0.0406
-6.250 -0.5614 0.02735 0.01895 -0.0300 1.0000 0.0426
-6.000 -0.5456 0.02598 0.01743 -0.0282 1.0000 0.0442
-5.750 -0.5294 0.02494 0.01622 -0.0264 1.0000 0.0467
-5.500 -0.5126 0.02397 0.01503 -0.0246 1.0000 0.0496
-5.250 -0.4787 0.02264 0.01341 -0.0261 0.9950 0.0517
-5.000 -0.4460 0.02134 0.01203 -0.0274 0.9892 0.0539
-4.750 -0.4126 0.02056 0.01120 -0.0290 0.9830 0.0581
-4.500 -0.3792 0.01973 0.01026 -0.0304 0.9766 0.0615
-4.250 -0.3476 0.01899 0.00937 -0.0313 0.9691 0.0643
-4.000 -0.3143 0.01821 0.00857 -0.0327 0.9629 0.0688
-3.750 -0.2835 0.01764 0.00790 -0.0335 0.9544 0.0759
-3.500 -0.2508 0.01699 0.00726 -0.0347 0.9474 0.0902
-3.250 -0.2195 0.01624 0.00669 -0.0357 0.9395 0.1283
-3.000 -0.1884 0.01539 0.00631 -0.0368 0.9322 0.2205
-2.750 -0.1566 0.01480 0.00605 -0.0379 0.9245 0.3181
-2.500 -0.1258 0.01425 0.00582 -0.0387 0.9164 0.4072
-2.250 -0.0935 0.01376 0.00561 -0.0395 0.9086 0.4929
-2.000 -0.0643 0.01320 0.00554 -0.0394 0.8998 0.6146
-1.750 -0.0291 0.01271 0.00555 -0.0398 0.8933 0.7710
-1.500 0.0390 0.01259 0.00562 -0.0466 0.8911 0.8959
-1.250 0.0902 0.01259 0.00548 -0.0510 0.8830 0.9356
-1.000 0.1408 0.01254 0.00529 -0.0554 0.8752 0.9589
-0.750 0.1884 0.01249 0.00510 -0.0594 0.8656 0.9774
-0.500 0.2366 0.01242 0.00491 -0.0637 0.8552 0.9928
-0.250 0.2758 0.01237 0.00475 -0.0662 0.8437 1.0000
0.000 0.3029 0.01237 0.00464 -0.0660 0.8308 1.0000
0.250 0.3290 0.01240 0.00458 -0.0657 0.8179 1.0000
0.500 0.3545 0.01245 0.00456 -0.0652 0.8053 1.0000
0.750 0.3785 0.01252 0.00458 -0.0644 0.7920 1.0000
1.000 0.4017 0.01261 0.00462 -0.0635 0.7785 1.0000
1.250 0.4248 0.01271 0.00470 -0.0625 0.7651 1.0000
1.500 0.4478 0.01281 0.00478 -0.0614 0.7517 1.0000
1.750 0.4710 0.01292 0.00487 -0.0604 0.7385 1.0000
2.000 0.4942 0.01303 0.00499 -0.0594 0.7252 1.0000
2.250 0.5175 0.01314 0.00510 -0.0584 0.7120 1.0000
2.500 0.5409 0.01325 0.00521 -0.0574 0.6987 1.0000
2.750 0.5644 0.01337 0.00535 -0.0563 0.6852 1.0000
3.000 0.5876 0.01349 0.00547 -0.0552 0.6697 1.0000
3.250 0.6104 0.01361 0.00557 -0.0539 0.6516 1.0000
3.500 0.6331 0.01373 0.00564 -0.0525 0.6315 1.0000
3.750 0.6547 0.01390 0.00581 -0.0510 0.6087 1.0000
4.000 0.6768 0.01410 0.00596 -0.0496 0.5871 1.0000
4.250 0.6986 0.01431 0.00621 -0.0482 0.5666 1.0000
4.500 0.7211 0.01454 0.00652 -0.0470 0.5492 1.0000
4.750 0.7429 0.01477 0.00678 -0.0456 0.5283 1.0000
5.000 0.7625 0.01502 0.00701 -0.0438 0.4951 1.0000
5.250 0.7807 0.01535 0.00723 -0.0418 0.4522 1.0000
5.500 0.7993 0.01575 0.00757 -0.0399 0.4112 1.0000
5.750 0.8168 0.01626 0.00795 -0.0379 0.3599 1.0000
6.000 0.8305 0.01706 0.00842 -0.0355 0.2850 1.0000
6.250 0.8405 0.01830 0.00913 -0.0328 0.2004 1.0000
6.500 0.8528 0.01951 0.00998 -0.0305 0.1401 1.0000
6.750 0.8652 0.02074 0.01094 -0.0283 0.0921 1.0000
7.000 0.8786 0.02190 0.01195 -0.0261 0.0626 1.0000
7.250 0.8920 0.02303 0.01302 -0.0239 0.0516 1.0000
7.500 0.9034 0.02430 0.01429 -0.0214 0.0451 1.0000
7.750 0.9166 0.02541 0.01554 -0.0192 0.0413 1.0000
8.000 0.9288 0.02660 0.01681 -0.0169 0.0373 1.0000
8.250 0.9364 0.02825 0.01847 -0.0143 0.0343 1.0000
8.500 0.9505 0.02953 0.01990 -0.0122 0.0327 1.0000
8.750 0.9652 0.03098 0.02149 -0.0104 0.0307 1.0000
9.000 0.9797 0.03244 0.02310 -0.0087 0.0284 1.0000
9.250 0.9935 0.03404 0.02476 -0.0072 0.0265 1.0000
9.500 1.0111 0.03675 0.02750 -0.0064 0.0250 1.0000
9.750 1.0275 0.03874 0.02975 -0.0050 0.0244 1.0000
10.000 1.0407 0.04086 0.03219 -0.0033 0.0236 1.0000
10.250 1.0492 0.04298 0.03463 -0.0011 0.0226 1.0000
10.500 1.0530 0.04511 0.03705 0.0014 0.0214 1.0000
10.750 1.0543 0.04749 0.03972 0.0040 0.0208 1.0000
11.000 1.0524 0.04994 0.04243 0.0064 0.0202 1.0000
11.250 1.0472 0.05279 0.04556 0.0086 0.0200 1.0000
11.500 1.0387 0.05588 0.04891 0.0104 0.0198 1.0000
11.750 1.0277 0.05927 0.05257 0.0117 0.0197 1.0000
12.000 1.0139 0.06309 0.05663 0.0122 0.0196 1.0000
12.250 0.9978 0.06741 0.06119 0.0120 0.0195 1.0000
12.500 0.9793 0.07240 0.06642 0.0107 0.0196 1.0000
12.750 0.9580 0.07832 0.07257 0.0082 0.0197 1.0000
13.000 0.9353 0.08510 0.07954 0.0046 0.0198 1.0000
13.250 0.9094 0.09341 0.08804 -0.0006 0.0202 1.0000
13.500 0.8789 0.10408 0.09886 -0.0076 0.0206 1.0000
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Polar data table (+)
Polar graphs
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