USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 46 AIRFOIL (usa46-il) Reynolds number: 200,000 Max Cl/Cd: 68.97 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa46-il-200000.txt Download as CSV file: xf-usa46-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 46 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4852 0.08619 0.08270 -0.0145 1.0000 0.0510
-7.750 -0.4875 0.08273 0.07930 -0.0164 1.0000 0.0528
-7.500 -0.4904 0.07783 0.07446 -0.0247 1.0000 0.0552
-7.250 -0.4824 0.07074 0.06731 -0.0374 1.0000 0.0561
-7.000 -0.4782 0.06333 0.05980 -0.0431 1.0000 0.0566
-6.750 -0.4722 0.05935 0.05589 -0.0417 1.0000 0.0578
-6.500 -0.4701 0.04059 0.03657 -0.0518 1.0000 0.0350
-6.250 -0.4538 0.03529 0.03090 -0.0537 1.0000 0.0358
-6.000 -0.4360 0.02791 0.02264 -0.0556 1.0000 0.0355
-5.750 -0.4144 0.02407 0.01824 -0.0557 1.0000 0.0362
-5.500 -0.3914 0.02158 0.01534 -0.0552 1.0000 0.0382
-5.250 -0.3674 0.01967 0.01302 -0.0546 1.0000 0.0415
-5.000 -0.3438 0.01775 0.01076 -0.0540 1.0000 0.0464
-4.750 -0.3198 0.01666 0.00954 -0.0533 1.0000 0.0520
-4.500 -0.2958 0.01537 0.00803 -0.0525 1.0000 0.0595
-4.250 -0.2717 0.01486 0.00741 -0.0517 1.0000 0.0681
-4.000 -0.2483 0.01396 0.00652 -0.0510 1.0000 0.0770
-3.750 -0.2245 0.01327 0.00576 -0.0502 1.0000 0.0856
-3.500 -0.2008 0.01292 0.00534 -0.0494 1.0000 0.0953
-3.250 -0.1773 0.01237 0.00483 -0.0486 1.0000 0.1059
-3.000 -0.1539 0.01203 0.00451 -0.0478 1.0000 0.1202
-2.750 -0.1308 0.01177 0.00431 -0.0470 1.0000 0.1400
-2.500 -0.1078 0.01157 0.00410 -0.0462 1.0000 0.1643
-2.250 -0.0850 0.01128 0.00389 -0.0455 1.0000 0.1859
-2.000 -0.0625 0.01115 0.00378 -0.0447 1.0000 0.2068
-1.750 -0.0405 0.01096 0.00369 -0.0439 1.0000 0.2257
-1.500 -0.0157 0.01083 0.00364 -0.0438 0.9991 0.2469
-1.250 0.0289 0.01052 0.00355 -0.0476 0.9914 0.2876
-1.000 0.0715 0.00954 0.00359 -0.0511 0.9835 0.5219
-0.750 0.1036 0.00871 0.00370 -0.0511 0.9732 0.8121
-0.500 0.1620 0.00824 0.00342 -0.0569 0.9684 1.0000
-0.250 0.2080 0.00814 0.00322 -0.0605 0.9577 1.0000
0.000 0.2518 0.00803 0.00304 -0.0635 0.9458 1.0000
0.250 0.2923 0.00793 0.00290 -0.0658 0.9329 1.0000
0.500 0.3307 0.00786 0.00278 -0.0676 0.9184 1.0000
0.750 0.3667 0.00781 0.00269 -0.0688 0.9018 1.0000
1.000 0.3976 0.00781 0.00265 -0.0690 0.8810 1.0000
1.250 0.4265 0.00786 0.00265 -0.0686 0.8590 1.0000
1.500 0.4531 0.00794 0.00267 -0.0679 0.8349 1.0000
1.750 0.4793 0.00804 0.00271 -0.0671 0.8098 1.0000
2.000 0.5051 0.00816 0.00276 -0.0662 0.7834 1.0000
2.250 0.5306 0.00831 0.00284 -0.0653 0.7550 1.0000
2.500 0.5559 0.00848 0.00293 -0.0643 0.7253 1.0000
2.750 0.5811 0.00867 0.00304 -0.0634 0.6940 1.0000
3.000 0.6061 0.00891 0.00317 -0.0624 0.6615 1.0000
3.250 0.6307 0.00916 0.00335 -0.0614 0.6237 1.0000
3.500 0.6545 0.00949 0.00351 -0.0603 0.5788 1.0000
3.750 0.6772 0.00992 0.00368 -0.0590 0.5204 1.0000
4.000 0.7004 0.01033 0.00390 -0.0579 0.4578 1.0000
4.250 0.7230 0.01085 0.00416 -0.0569 0.3863 1.0000
4.500 0.7449 0.01157 0.00454 -0.0559 0.3166 1.0000
4.750 0.7591 0.01370 0.00540 -0.0543 0.0850 1.0000
5.000 0.7796 0.01527 0.00676 -0.0529 0.0450 1.0000
5.250 0.8014 0.01647 0.00804 -0.0517 0.0388 1.0000
5.500 0.8245 0.01735 0.00898 -0.0507 0.0340 1.0000
5.750 0.8454 0.01874 0.01041 -0.0495 0.0315 1.0000
6.000 0.8653 0.02114 0.01282 -0.0480 0.0300 1.0000
6.250 0.8895 0.02292 0.01469 -0.0470 0.0295 1.0000
6.500 0.9142 0.02466 0.01660 -0.0461 0.0290 1.0000
6.750 0.9384 0.02600 0.01817 -0.0451 0.0272 1.0000
7.000 0.9617 0.02840 0.02086 -0.0440 0.0270 1.0000
7.250 0.9830 0.03156 0.02443 -0.0426 0.0276 1.0000
7.500 1.0002 0.03611 0.02937 -0.0411 0.0291 1.0000
11.250 0.7625 0.12703 0.12403 -0.0577 0.0578 1.0000
11.500 0.7632 0.13104 0.12805 -0.0588 0.0562 1.0000
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