RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAF 28 AIRFOIL (raf28-il) Reynolds number: 200,000 Max Cl/Cd: 69.45 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-raf28-il-200000.txt Download as CSV file: xf-raf28-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4898 0.11569 0.11189 -0.0200 1.0000 0.0533
-10.750 -0.5023 0.11101 0.10727 -0.0262 1.0000 0.0548
-10.500 -0.5106 0.10618 0.10250 -0.0309 1.0000 0.0551
-10.250 -0.4990 0.10193 0.09825 -0.0268 1.0000 0.0567
-10.000 -0.4885 0.09967 0.09598 -0.0255 1.0000 0.0583
-9.750 -0.4843 0.09667 0.09299 -0.0259 1.0000 0.0605
-9.500 -0.4860 0.09280 0.08917 -0.0278 1.0000 0.0635
-9.250 -0.4042 0.07904 0.07566 -0.0330 1.0000 0.0709
-9.000 -0.4015 0.07596 0.07259 -0.0329 1.0000 0.0742
-8.750 -0.4110 0.07126 0.06793 -0.0349 1.0000 0.0768
-8.250 -0.5551 0.06659 0.06315 -0.0439 1.0000 0.0695
-8.000 -0.5479 0.06460 0.06118 -0.0422 1.0000 0.0710
-7.750 -0.5470 0.06183 0.05839 -0.0414 1.0000 0.0737
-7.500 -0.5555 0.05750 0.05390 -0.0417 1.0000 0.0775
-7.250 -0.5799 0.05213 0.04801 -0.0405 1.0000 0.0817
-7.000 -0.5620 0.05000 0.04615 -0.0390 1.0000 0.0844
-6.750 -0.5556 0.04786 0.04395 -0.0370 1.0000 0.0881
-6.500 -0.5823 0.03422 0.02890 -0.0312 1.0000 0.0524
-6.250 -0.5799 0.03024 0.02448 -0.0276 1.0000 0.0492
-6.000 -0.5732 0.02672 0.02036 -0.0242 1.0000 0.0474
-5.750 -0.5600 0.02526 0.01869 -0.0220 1.0000 0.0488
-5.500 -0.5422 0.02384 0.01699 -0.0205 0.9995 0.0506
-5.250 -0.5064 0.02180 0.01456 -0.0222 0.9956 0.0517
-5.000 -0.4702 0.02055 0.01300 -0.0239 0.9913 0.0540
-4.750 -0.4340 0.01917 0.01138 -0.0256 0.9869 0.0557
-4.500 -0.3968 0.01766 0.00983 -0.0277 0.9835 0.0581
-4.250 -0.3631 0.01690 0.00906 -0.0291 0.9773 0.0617
-4.000 -0.3247 0.01617 0.00826 -0.0313 0.9726 0.0650
-3.750 -0.2900 0.01542 0.00745 -0.0327 0.9666 0.0680
-3.500 -0.2547 0.01462 0.00672 -0.0344 0.9605 0.0737
-3.250 -0.2150 0.01397 0.00610 -0.0369 0.9562 0.0859
-3.000 -0.1851 0.01301 0.00541 -0.0375 0.9478 0.1375
-2.750 -0.1496 0.01175 0.00505 -0.0397 0.9434 0.3418
-2.500 -0.1198 0.01118 0.00485 -0.0402 0.9348 0.4375
-2.250 -0.0846 0.01044 0.00464 -0.0416 0.9299 0.5693
-2.000 -0.0579 0.00990 0.00458 -0.0409 0.9210 0.6947
-1.750 -0.0205 0.00947 0.00446 -0.0420 0.9157 0.7970
-1.500 0.0454 0.00934 0.00451 -0.0486 0.9146 0.9026
-1.250 0.1108 0.00948 0.00459 -0.0555 0.9110 0.9518
-1.000 0.1749 0.00956 0.00457 -0.0624 0.9054 0.9741
-0.750 0.2304 0.00946 0.00436 -0.0680 0.8977 0.9852
-0.500 0.2836 0.00932 0.00413 -0.0732 0.8867 0.9950
-0.250 0.3243 0.00920 0.00391 -0.0760 0.8730 1.0000
0.000 0.3479 0.00918 0.00379 -0.0752 0.8572 1.0000
0.250 0.3702 0.00921 0.00372 -0.0741 0.8419 1.0000
0.500 0.3920 0.00926 0.00369 -0.0729 0.8270 1.0000
0.750 0.4136 0.00933 0.00368 -0.0716 0.8124 1.0000
1.000 0.4353 0.00940 0.00370 -0.0703 0.7982 1.0000
1.250 0.4570 0.00949 0.00374 -0.0689 0.7841 1.0000
1.500 0.4788 0.00956 0.00377 -0.0676 0.7698 1.0000
1.750 0.5006 0.00962 0.00380 -0.0662 0.7552 1.0000
2.000 0.5226 0.00967 0.00381 -0.0649 0.7399 1.0000
2.250 0.5447 0.00970 0.00380 -0.0634 0.7236 1.0000
2.500 0.5667 0.00974 0.00380 -0.0620 0.7064 1.0000
2.750 0.5886 0.00980 0.00383 -0.0606 0.6884 1.0000
3.000 0.6109 0.00989 0.00389 -0.0592 0.6704 1.0000
3.250 0.6333 0.01000 0.00395 -0.0579 0.6526 1.0000
3.500 0.6559 0.01016 0.00404 -0.0566 0.6349 1.0000
3.750 0.6776 0.01031 0.00418 -0.0552 0.6141 1.0000
4.000 0.6991 0.01049 0.00430 -0.0537 0.5918 1.0000
4.250 0.7207 0.01066 0.00448 -0.0522 0.5701 1.0000
4.500 0.7421 0.01084 0.00464 -0.0508 0.5475 1.0000
4.750 0.7623 0.01103 0.00479 -0.0491 0.5181 1.0000
5.000 0.7820 0.01126 0.00498 -0.0473 0.4834 1.0000
5.250 0.8003 0.01158 0.00518 -0.0453 0.4408 1.0000
5.500 0.8161 0.01207 0.00545 -0.0429 0.3793 1.0000
5.750 0.8266 0.01298 0.00586 -0.0398 0.2779 1.0000
6.000 0.8319 0.01454 0.00667 -0.0362 0.1546 1.0000
6.250 0.8396 0.01609 0.00767 -0.0329 0.0754 1.0000
6.500 0.8545 0.01703 0.00851 -0.0306 0.0587 1.0000
6.750 0.8673 0.01813 0.00958 -0.0280 0.0516 1.0000
7.000 0.8831 0.01899 0.01054 -0.0258 0.0481 1.0000
7.250 0.8975 0.01998 0.01156 -0.0236 0.0442 1.0000
7.500 0.9073 0.02171 0.01324 -0.0208 0.0409 1.0000
7.750 0.9254 0.02273 0.01438 -0.0191 0.0395 1.0000
8.000 0.9445 0.02399 0.01573 -0.0175 0.0380 1.0000
8.250 0.9651 0.02543 0.01724 -0.0163 0.0363 1.0000
8.500 0.9855 0.02685 0.01872 -0.0153 0.0344 1.0000
8.750 1.0083 0.02923 0.02110 -0.0149 0.0323 1.0000
9.000 1.0324 0.03270 0.02481 -0.0146 0.0317 1.0000
9.250 1.0509 0.03625 0.02868 -0.0135 0.0315 1.0000
9.500 1.0660 0.03898 0.03170 -0.0118 0.0315 1.0000
9.750 1.0769 0.04211 0.03515 -0.0098 0.0315 1.0000
10.000 1.0799 0.04317 0.03658 -0.0063 0.0308 1.0000
10.250 1.0827 0.04623 0.03996 -0.0036 0.0308 1.0000
10.500 1.0820 0.04944 0.04347 -0.0007 0.0308 1.0000
10.750 1.0959 0.05183 0.04605 0.0009 0.0324 1.0000
11.000 1.0443 0.05877 0.05385 0.0090 0.0393 1.0000
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Polar data table (+)
Polar graphs
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