ARA-D 20% AIRFOIL (arad20-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: ARA-D 20% AIRFOIL (arad20-il) Reynolds number: 50,000 Max Cl/Cd: 9.72 at α=-0.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-arad20-il-50000.txt Download as CSV file: xf-arad20-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: ARA-D 20% AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3083 0.12117 0.11521 0.0066 1.0000 0.3397
-9.500 -0.5388 0.08005 0.07378 -0.0304 1.0000 0.1713
-9.250 -0.5550 0.07396 0.06766 -0.0316 1.0000 0.1700
-9.000 -0.5773 0.06755 0.06112 -0.0327 1.0000 0.1688
-8.750 -0.5954 0.06152 0.05484 -0.0332 1.0000 0.1686
-8.500 -0.6059 0.05639 0.04939 -0.0333 1.0000 0.1696
-8.250 -0.6352 0.05652 0.04971 -0.0300 0.9960 0.1691
-8.000 -0.5847 0.05127 0.04410 -0.0382 0.9729 0.1767
-7.750 -0.5388 0.04626 0.03829 -0.0454 0.9516 0.1863
-7.500 -0.4848 0.04348 0.03549 -0.0509 0.9286 0.1983
-7.250 -0.4456 0.04103 0.03281 -0.0538 0.9042 0.2100
-7.000 -0.4149 0.03900 0.03041 -0.0548 0.8810 0.2227
-6.750 -0.3869 0.03767 0.02884 -0.0548 0.8601 0.2370
-6.500 -0.3605 0.03661 0.02780 -0.0540 0.8420 0.2525
-6.250 -0.3355 0.03566 0.02682 -0.0534 0.8240 0.2702
-6.000 -0.3104 0.03482 0.02594 -0.0528 0.8083 0.2917
-5.750 -0.2858 0.03404 0.02511 -0.0516 0.7953 0.3171
-5.500 -0.2604 0.03359 0.02491 -0.0510 0.7808 0.3453
-5.250 -0.2358 0.03311 0.02447 -0.0499 0.7694 0.3806
-5.000 -0.2111 0.03288 0.02443 -0.0490 0.7567 0.4196
-4.750 -0.1874 0.03283 0.02458 -0.0471 0.7466 0.4609
-4.500 -0.1632 0.03301 0.02499 -0.0457 0.7349 0.5043
-4.250 -0.1412 0.03328 0.02538 -0.0426 0.7260 0.5471
-4.000 -0.1183 0.03389 0.02618 -0.0409 0.7148 0.5878
-3.750 -0.0971 0.03438 0.02671 -0.0377 0.7058 0.6276
-3.500 -0.0765 0.03511 0.02747 -0.0350 0.6969 0.6657
-3.250 -0.0566 0.03592 0.02835 -0.0317 0.6871 0.6984
-3.000 -0.0371 0.03648 0.02882 -0.0276 0.6796 0.7330
-2.750 -0.0175 0.03745 0.02982 -0.0255 0.6703 0.7647
-2.500 0.0049 0.03820 0.03053 -0.0225 0.6614 0.7952
-2.250 0.0275 0.03852 0.03067 -0.0192 0.6548 0.8295
-2.000 0.0609 0.03992 0.03210 -0.0204 0.6439 0.8575
-1.750 0.1077 0.04051 0.03251 -0.0220 0.6347 0.8881
-1.500 0.1661 0.04176 0.03363 -0.0278 0.6239 0.9159
-1.250 0.2487 0.04260 0.03424 -0.0374 0.6118 0.9404
-1.000 0.3155 0.04372 0.03523 -0.0461 0.6008 0.9650
-0.750 0.3903 0.04444 0.03579 -0.0561 0.5892 0.9868
-0.500 0.4388 0.04514 0.03635 -0.0619 0.5808 1.0000
-0.250 0.4419 0.04684 0.03812 -0.0627 0.5729 1.0000
0.000 0.4465 0.04772 0.03893 -0.0616 0.5670 1.0000
0.250 0.4545 0.04803 0.03908 -0.0594 0.5622 1.0000
0.500 0.4429 0.05070 0.04183 -0.0586 0.5562 1.0000
0.750 0.4154 0.05396 0.04517 -0.0565 0.5509 1.0000
1.000 0.3811 0.05701 0.04824 -0.0529 0.5479 1.0000
1.250 0.3399 0.06047 0.05168 -0.0489 0.5467 1.0000
1.500 0.3014 0.06473 0.05590 -0.0465 0.5483 1.0000
1.750 0.2843 0.06873 0.05985 -0.0461 0.5526 1.0000
2.000 0.2902 0.07213 0.06317 -0.0470 0.5567 1.0000
2.250 0.3107 0.07526 0.06621 -0.0486 0.5594 1.0000
2.500 0.1893 0.08545 0.07676 -0.0524 0.6816 1.0000
2.750 0.1912 0.08688 0.07809 -0.0510 0.6654 1.0000
3.000 0.2407 0.09099 0.08205 -0.0547 0.6558 1.0000
3.250 0.2295 0.09143 0.08244 -0.0519 0.6386 1.0000
3.500 0.2798 0.09610 0.08697 -0.0556 0.6318 1.0000
3.750 0.2610 0.09607 0.08689 -0.0523 0.6150 1.0000
4.000 0.3136 0.10103 0.09173 -0.0559 0.6081 1.0000
4.250 0.2896 0.10076 0.09142 -0.0526 0.5921 1.0000
4.500 0.3328 0.10504 0.09560 -0.0551 0.5854 1.0000
4.750 0.3134 0.10563 0.09617 -0.0527 0.5726 1.0000
5.000 0.3467 0.10898 0.09944 -0.0542 0.5637 1.0000
5.250 0.3472 0.11136 0.10177 -0.0538 0.5558 1.0000
5.500 0.3578 0.11323 0.10358 -0.0536 0.5441 1.0000
5.750 0.4003 0.11802 0.10830 -0.0559 0.5387 1.0000
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