NACA 2.5411 (naca2411-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 2.5411 (naca2411-il) Reynolds number: 200,000 Max Cl/Cd: 71.21 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca2411-il-200000.txt Download as CSV file: xf-naca2411-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 2.5411
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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4670 0.08689 0.08323 -0.0431 1.0000 0.0774
-9.250 -0.4585 0.08487 0.08122 -0.0412 1.0000 0.0784
-9.000 -0.6182 0.05844 0.05460 -0.0544 1.0000 0.0552
-8.750 -0.6049 0.05764 0.05392 -0.0522 1.0000 0.0559
-8.500 -0.6866 0.04320 0.03837 -0.0472 1.0000 0.0454
-8.250 -0.6905 0.03949 0.03436 -0.0443 1.0000 0.0452
-8.000 -0.6901 0.03621 0.03073 -0.0416 1.0000 0.0451
-7.750 -0.6854 0.03380 0.02787 -0.0389 1.0000 0.0455
-7.500 -0.6790 0.03043 0.02423 -0.0368 1.0000 0.0465
-7.250 -0.6655 0.02892 0.02263 -0.0351 1.0000 0.0477
-7.000 -0.6509 0.02747 0.02101 -0.0334 1.0000 0.0486
-6.750 -0.6354 0.02595 0.01926 -0.0317 1.0000 0.0497
-6.500 -0.6185 0.02456 0.01761 -0.0301 1.0000 0.0508
-6.250 -0.5914 0.02321 0.01595 -0.0303 0.9981 0.0527
-6.000 -0.5554 0.02174 0.01422 -0.0323 0.9943 0.0559
-5.750 -0.5200 0.02088 0.01334 -0.0342 0.9896 0.0593
-5.500 -0.4816 0.02005 0.01232 -0.0364 0.9852 0.0634
-5.250 -0.4464 0.01902 0.01129 -0.0382 0.9803 0.0687
-5.000 -0.4089 0.01855 0.01068 -0.0401 0.9747 0.0756
-4.750 -0.3691 0.01769 0.00990 -0.0428 0.9711 0.0842
-4.500 -0.3383 0.01706 0.00928 -0.0435 0.9640 0.0936
-4.250 -0.2999 0.01651 0.00874 -0.0457 0.9594 0.1064
-4.000 -0.2642 0.01600 0.00829 -0.0474 0.9541 0.1212
-3.750 -0.2309 0.01552 0.00788 -0.0486 0.9474 0.1391
-3.500 -0.1907 0.01499 0.00745 -0.0511 0.9435 0.1654
-3.250 -0.1601 0.01456 0.00715 -0.0517 0.9361 0.1947
-3.000 -0.1230 0.01407 0.00684 -0.0536 0.9308 0.2344
-2.750 -0.0811 0.01356 0.00652 -0.0565 0.9277 0.2851
-2.500 -0.0534 0.01315 0.00631 -0.0564 0.9188 0.3338
-2.250 -0.0143 0.01253 0.00600 -0.0586 0.9145 0.4048
-2.000 0.0241 0.01180 0.00575 -0.0606 0.9114 0.5156
-1.750 0.0449 0.01126 0.00566 -0.0587 0.9007 0.6297
-1.500 0.0801 0.01063 0.00551 -0.0591 0.8965 0.7581
-1.250 0.1178 0.01038 0.00552 -0.0600 0.8886 0.8615
-1.000 0.1821 0.01027 0.00543 -0.0664 0.8852 0.9260
-0.750 0.2429 0.01017 0.00524 -0.0724 0.8810 0.9557
-0.500 0.3002 0.01012 0.00512 -0.0780 0.8720 0.9738
-0.250 0.3616 0.00995 0.00485 -0.0844 0.8645 0.9878
0.000 0.4174 0.00977 0.00460 -0.0901 0.8516 1.0000
0.250 0.4404 0.00966 0.00439 -0.0890 0.8352 1.0000
0.500 0.4632 0.00959 0.00422 -0.0878 0.8184 1.0000
0.750 0.4859 0.00955 0.00408 -0.0866 0.8014 1.0000
1.000 0.5071 0.00956 0.00401 -0.0850 0.7829 1.0000
1.250 0.5287 0.00960 0.00396 -0.0836 0.7646 1.0000
1.500 0.5506 0.00966 0.00393 -0.0822 0.7466 1.0000
1.750 0.5726 0.00975 0.00393 -0.0808 0.7289 1.0000
2.000 0.5947 0.00986 0.00395 -0.0795 0.7115 1.0000
2.250 0.6168 0.00999 0.00399 -0.0781 0.6942 1.0000
2.500 0.6390 0.01013 0.00405 -0.0768 0.6774 1.0000
2.750 0.6611 0.01030 0.00413 -0.0754 0.6608 1.0000
3.000 0.6831 0.01047 0.00423 -0.0741 0.6444 1.0000
3.250 0.7051 0.01066 0.00435 -0.0727 0.6280 1.0000
3.500 0.7269 0.01085 0.00448 -0.0713 0.6117 1.0000
3.750 0.7485 0.01104 0.00464 -0.0699 0.5950 1.0000
4.000 0.7699 0.01123 0.00479 -0.0685 0.5778 1.0000
4.250 0.7910 0.01143 0.00496 -0.0670 0.5603 1.0000
4.500 0.8120 0.01164 0.00514 -0.0654 0.5422 1.0000
4.750 0.8326 0.01185 0.00531 -0.0638 0.5237 1.0000
5.000 0.8530 0.01208 0.00548 -0.0622 0.5046 1.0000
5.250 0.8733 0.01234 0.00567 -0.0605 0.4855 1.0000
5.500 0.8932 0.01256 0.00591 -0.0589 0.4649 1.0000
5.750 0.9129 0.01282 0.00614 -0.0572 0.4438 1.0000
6.000 0.9319 0.01310 0.00638 -0.0554 0.4210 1.0000
6.250 0.9503 0.01339 0.00663 -0.0535 0.3946 1.0000
6.500 0.9678 0.01374 0.00693 -0.0514 0.3647 1.0000
6.750 0.9838 0.01419 0.00725 -0.0492 0.3307 1.0000
7.000 0.9987 0.01473 0.00765 -0.0469 0.2906 1.0000
7.250 1.0113 0.01545 0.00815 -0.0443 0.2459 1.0000
7.500 1.0220 0.01634 0.00878 -0.0415 0.1989 1.0000
7.750 1.0315 0.01737 0.00953 -0.0385 0.1565 1.0000
8.000 1.0407 0.01842 0.01038 -0.0356 0.1266 1.0000
8.250 1.0502 0.01944 0.01126 -0.0327 0.1071 1.0000
8.500 1.0591 0.02040 0.01215 -0.0297 0.0937 1.0000
8.750 1.0653 0.02145 0.01313 -0.0263 0.0841 1.0000
9.000 1.0769 0.02225 0.01399 -0.0237 0.0763 1.0000
9.250 1.0835 0.02343 0.01514 -0.0206 0.0703 1.0000
9.500 1.0947 0.02434 0.01609 -0.0183 0.0650 1.0000
9.750 1.1015 0.02575 0.01744 -0.0156 0.0606 1.0000
10.000 1.1140 0.02669 0.01851 -0.0135 0.0568 1.0000
10.250 1.1247 0.02785 0.01967 -0.0115 0.0536 1.0000
10.500 1.1377 0.02959 0.02136 -0.0099 0.0510 1.0000
10.750 1.1506 0.03068 0.02261 -0.0081 0.0485 1.0000
11.000 1.1636 0.03188 0.02388 -0.0065 0.0463 1.0000
11.250 1.1766 0.03319 0.02518 -0.0051 0.0443 1.0000
11.500 1.1976 0.03534 0.02737 -0.0046 0.0424 1.0000
11.750 1.2097 0.03682 0.02904 -0.0030 0.0413 1.0000
12.000 1.2225 0.03854 0.03095 -0.0016 0.0403 1.0000
12.250 1.2333 0.04030 0.03288 -0.0001 0.0392 1.0000
12.500 1.2423 0.04208 0.03481 0.0015 0.0382 1.0000
12.750 1.2501 0.04390 0.03674 0.0029 0.0373 1.0000
13.000 1.2612 0.04580 0.03865 0.0038 0.0362 1.0000
13.250 1.2702 0.04949 0.04252 0.0046 0.0354 1.0000
13.500 1.2639 0.05201 0.04530 0.0067 0.0351 1.0000
13.750 1.2564 0.05487 0.04843 0.0084 0.0350 1.0000
14.000 1.2445 0.05811 0.05193 0.0098 0.0348 1.0000
14.250 1.2314 0.06181 0.05589 0.0106 0.0347 1.0000
14.500 1.2170 0.06593 0.06026 0.0109 0.0347 1.0000
14.750 1.1987 0.07059 0.06516 0.0104 0.0346 1.0000
15.000 1.1791 0.07573 0.07052 0.0093 0.0347 1.0000
15.250 1.1578 0.08153 0.07654 0.0074 0.0348 1.0000
15.500 1.1347 0.08791 0.08313 0.0047 0.0348 1.0000
15.750 1.1109 0.09501 0.09041 0.0012 0.0350 1.0000
16.000 1.0844 0.10310 0.09868 -0.0032 0.0352 1.0000
16.250 0.6913 0.16945 0.16608 -0.0420 0.0720 1.0000
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Polar data table (+)
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