RAE 5215 AIRFOIL (rae5215-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 5215 AIRFOIL (rae5215-il) Reynolds number: 200,000 Max Cl/Cd: 43.76 at α=2° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae5215-il-200000.txt Download as CSV file: xf-rae5215-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 5215 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4266 0.11178 0.10812 -0.0153 1.0000 0.0525
-10.500 -0.4358 0.10679 0.10314 -0.0184 1.0000 0.0544
-10.250 -0.4653 0.09892 0.09531 -0.0259 1.0000 0.0552
-10.000 -0.4515 0.09597 0.09238 -0.0233 1.0000 0.0559
-9.750 -0.4404 0.09356 0.08998 -0.0220 1.0000 0.0569
-9.500 -0.4374 0.09002 0.08645 -0.0224 1.0000 0.0582
-9.250 -0.4398 0.08560 0.08206 -0.0238 1.0000 0.0598
-9.000 -0.4516 0.07954 0.07603 -0.0273 1.0000 0.0617
-8.750 -0.5744 0.08048 0.07683 -0.0284 1.0000 0.0567
-8.500 -0.5764 0.07582 0.07218 -0.0317 1.0000 0.0575
-8.250 -0.5860 0.07002 0.06636 -0.0377 1.0000 0.0582
-8.000 -0.5899 0.06505 0.06129 -0.0416 1.0000 0.0596
-7.750 -0.6165 0.05831 0.05373 -0.0480 1.0000 0.0636
-7.500 -0.5995 0.05431 0.04997 -0.0474 1.0000 0.0647
-7.250 -0.5864 0.05168 0.04738 -0.0468 1.0000 0.0661
-7.000 -0.5755 0.04897 0.04455 -0.0465 1.0000 0.0688
-6.750 -0.5697 0.04503 0.04017 -0.0468 1.0000 0.0743
-6.500 -0.5576 0.03256 0.02608 -0.0457 1.0000 0.0425
-6.250 -0.5388 0.02996 0.02333 -0.0446 1.0000 0.0413
-6.000 -0.5194 0.02755 0.02047 -0.0435 1.0000 0.0417
-5.750 -0.4989 0.02553 0.01809 -0.0424 1.0000 0.0419
-5.500 -0.4776 0.02369 0.01596 -0.0413 1.0000 0.0421
-5.250 -0.4569 0.02203 0.01441 -0.0405 1.0000 0.0438
-5.000 -0.4350 0.02103 0.01331 -0.0396 1.0000 0.0458
-4.750 -0.4117 0.01981 0.01190 -0.0386 1.0000 0.0468
-4.500 -0.3880 0.01891 0.01079 -0.0377 1.0000 0.0481
-4.250 -0.3641 0.01772 0.00970 -0.0373 1.0000 0.0501
-4.000 -0.3404 0.01715 0.00914 -0.0367 1.0000 0.0534
-3.750 -0.3152 0.01640 0.00837 -0.0364 1.0000 0.0562
-3.500 -0.2895 0.01580 0.00786 -0.0364 1.0000 0.0601
-3.250 -0.2631 0.01530 0.00734 -0.0363 1.0000 0.0637
-3.000 -0.2355 0.01465 0.00679 -0.0368 1.0000 0.0690
-2.750 -0.2075 0.01415 0.00634 -0.0373 1.0000 0.0778
-2.500 -0.1774 0.01350 0.00585 -0.0382 1.0000 0.1051
-2.250 -0.1372 0.01147 0.00591 -0.0420 1.0000 0.5920
-2.000 -0.1048 0.01175 0.00621 -0.0428 0.9973 0.6349
-1.750 -0.0682 0.01204 0.00651 -0.0443 0.9940 0.6634
-1.500 -0.0361 0.01238 0.00692 -0.0447 0.9892 0.6961
-1.250 -0.0013 0.01274 0.00740 -0.0453 0.9851 0.7293
-1.000 0.0345 0.01282 0.00753 -0.0464 0.9775 0.7483
-0.750 0.0785 0.01271 0.00742 -0.0493 0.9701 0.7569
-0.500 0.1234 0.01254 0.00721 -0.0526 0.9624 0.7636
-0.250 0.1727 0.01215 0.00684 -0.0562 0.9513 0.7689
0.000 0.2214 0.01162 0.00630 -0.0593 0.9363 0.7751
0.250 0.2572 0.01125 0.00591 -0.0599 0.9207 0.7811
0.500 0.2863 0.01096 0.00566 -0.0591 0.9061 0.7864
0.750 0.3147 0.01077 0.00546 -0.0584 0.8925 0.7929
1.000 0.3414 0.01059 0.00528 -0.0575 0.8767 0.7988
1.250 0.3661 0.01042 0.00513 -0.0560 0.8571 0.8047
1.500 0.3914 0.01028 0.00499 -0.0547 0.8310 0.8115
1.750 0.4159 0.01012 0.00480 -0.0531 0.7976 0.8175
2.000 0.4393 0.01004 0.00459 -0.0512 0.7379 0.8244
2.250 0.4526 0.01114 0.00435 -0.0474 0.4519 0.8314
2.500 0.4679 0.01278 0.00486 -0.0457 0.2373 0.8383
2.750 0.4925 0.01338 0.00519 -0.0452 0.2016 0.8470
3.000 0.5165 0.01378 0.00550 -0.0444 0.1845 0.8552
3.250 0.5423 0.01412 0.00580 -0.0441 0.1723 0.8651
3.500 0.5661 0.01445 0.00615 -0.0431 0.1630 0.8761
3.750 0.5898 0.01482 0.00649 -0.0423 0.1548 0.8896
4.000 0.6134 0.01517 0.00688 -0.0413 0.1475 0.9077
4.250 0.6377 0.01546 0.00720 -0.0404 0.1405 0.9403
4.500 0.6692 0.01610 0.00781 -0.0415 0.1332 1.0000
4.750 0.6997 0.01663 0.00828 -0.0424 0.1257 1.0000
5.000 0.7287 0.01742 0.00902 -0.0430 0.1185 1.0000
5.250 0.7573 0.01789 0.00945 -0.0434 0.1114 1.0000
5.500 0.7849 0.01871 0.01023 -0.0436 0.1050 1.0000
5.750 0.8125 0.01915 0.01069 -0.0437 0.0987 1.0000
6.000 0.8392 0.02000 0.01149 -0.0437 0.0932 1.0000
6.250 0.8660 0.02044 0.01199 -0.0436 0.0875 1.0000
6.500 0.8920 0.02128 0.01278 -0.0435 0.0826 1.0000
6.750 0.9181 0.02175 0.01333 -0.0432 0.0774 1.0000
7.000 0.9436 0.02267 0.01418 -0.0430 0.0736 1.0000
7.250 0.9691 0.02346 0.01515 -0.0426 0.0695 1.0000
7.500 0.9944 0.02420 0.01578 -0.0424 0.0666 1.0000
7.750 1.0187 0.02541 0.01722 -0.0418 0.0634 1.0000
8.000 1.0430 0.02626 0.01815 -0.0414 0.0608 1.0000
8.250 1.0675 0.02735 0.01916 -0.0411 0.0589 1.0000
8.500 1.0894 0.02922 0.02136 -0.0402 0.0572 1.0000
8.750 1.1102 0.03114 0.02361 -0.0393 0.0557 1.0000
9.000 1.1310 0.03280 0.02546 -0.0385 0.0544 1.0000
9.250 1.1522 0.03409 0.02682 -0.0378 0.0532 1.0000
9.500 1.1728 0.03588 0.02862 -0.0373 0.0521 1.0000
9.750 1.1825 0.03908 0.03238 -0.0353 0.0513 1.0000
10.000 1.1867 0.04317 0.03705 -0.0330 0.0508 1.0000
10.250 1.1833 0.04809 0.04252 -0.0305 0.0506 1.0000
10.500 1.1720 0.05357 0.04847 -0.0277 0.0507 1.0000
10.750 1.1553 0.05900 0.05428 -0.0251 0.0510 1.0000
11.000 1.1353 0.06377 0.05931 -0.0225 0.0513 1.0000
11.250 1.1136 0.06818 0.06390 -0.0205 0.0515 1.0000
11.500 1.0938 0.07290 0.06875 -0.0200 0.0517 1.0000
11.750 1.0792 0.07775 0.07369 -0.0205 0.0520 1.0000
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Polar data table (+)
Polar graphs
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