USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 46 AIRFOIL (usa46-il) Reynolds number: 100,000 Max Cl/Cd: 52.92 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa46-il-100000.txt Download as CSV file: xf-usa46-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4911 0.10620 0.10110 -0.0062 1.0000 0.0921
-8.750 -0.5016 0.10424 0.09926 -0.0110 1.0000 0.0948
-8.500 -0.5152 0.10232 0.09746 -0.0159 1.0000 0.0954
-8.250 -0.4858 0.09569 0.09075 -0.0106 1.0000 0.0990
-8.000 -0.4798 0.09256 0.08764 -0.0109 1.0000 0.1028
-7.750 -0.4864 0.08994 0.08512 -0.0137 1.0000 0.1070
-7.500 -0.5009 0.08652 0.08178 -0.0299 1.0000 0.1088
-7.250 -0.4825 0.08221 0.07751 -0.0178 1.0000 0.1116
-7.000 -0.4734 0.07894 0.07426 -0.0178 1.0000 0.1160
-6.750 -0.4053 0.06351 0.05914 -0.0191 1.0000 0.1315
-6.500 -0.4613 0.07030 0.06567 -0.0267 1.0000 0.1258
-6.250 -0.4498 0.06523 0.06034 -0.0397 1.0000 0.1364
-6.000 -0.4393 0.06269 0.05800 -0.0311 1.0000 0.1414
-5.750 -0.4124 0.04390 0.03806 -0.0513 1.0000 0.0819
-5.500 -0.3946 0.03925 0.03334 -0.0517 1.0000 0.0793
-5.250 -0.3725 0.03307 0.02656 -0.0535 1.0000 0.0761
-5.000 -0.3483 0.02852 0.02129 -0.0542 1.0000 0.0790
-4.750 -0.3244 0.02549 0.01772 -0.0542 1.0000 0.0857
-4.500 -0.2995 0.02338 0.01519 -0.0537 1.0000 0.0931
-4.250 -0.2757 0.02214 0.01380 -0.0531 1.0000 0.1039
-4.000 -0.2511 0.02067 0.01205 -0.0525 1.0000 0.1136
-3.750 -0.2264 0.01953 0.01066 -0.0518 1.0000 0.1246
-3.500 -0.2016 0.01853 0.00944 -0.0510 1.0000 0.1354
-3.250 -0.1778 0.01774 0.00857 -0.0503 1.0000 0.1509
-3.000 -0.1539 0.01679 0.00764 -0.0495 1.0000 0.1670
-2.750 -0.1302 0.01595 0.00683 -0.0488 1.0000 0.1884
-2.500 -0.1068 0.01523 0.00619 -0.0480 1.0000 0.2182
-2.250 -0.0836 0.01455 0.00558 -0.0473 1.0000 0.2491
-2.000 -0.0610 0.01396 0.00521 -0.0465 1.0000 0.2814
-1.750 -0.0380 0.01348 0.00492 -0.0458 1.0000 0.3177
-1.500 -0.0157 0.01297 0.00472 -0.0450 1.0000 0.3690
-1.250 0.0051 0.01179 0.00481 -0.0435 1.0000 0.6250
-1.000 0.0304 0.01092 0.00454 -0.0418 1.0000 1.0000
-0.750 0.0532 0.01110 0.00449 -0.0413 1.0000 1.0000
-0.500 0.0754 0.01132 0.00454 -0.0408 1.0000 1.0000
-0.250 0.0968 0.01158 0.00467 -0.0403 1.0000 1.0000
0.000 0.1176 0.01190 0.00487 -0.0398 1.0000 1.0000
0.250 0.1376 0.01227 0.00516 -0.0393 1.0000 1.0000
0.500 0.1999 0.01267 0.00548 -0.0467 0.9835 1.0000
0.750 0.2549 0.01287 0.00565 -0.0524 0.9673 1.0000
1.000 0.3063 0.01295 0.00574 -0.0571 0.9513 1.0000
1.250 0.3577 0.01293 0.00575 -0.0616 0.9358 1.0000
1.500 0.4086 0.01282 0.00573 -0.0658 0.9202 1.0000
1.750 0.4553 0.01268 0.00566 -0.0688 0.9031 1.0000
2.000 0.4953 0.01258 0.00563 -0.0704 0.8826 1.0000
2.250 0.5304 0.01251 0.00565 -0.0709 0.8597 1.0000
2.500 0.5601 0.01253 0.00572 -0.0703 0.8335 1.0000
2.750 0.5876 0.01258 0.00580 -0.0692 0.8059 1.0000
3.000 0.6137 0.01267 0.00591 -0.0678 0.7766 1.0000
3.250 0.6391 0.01280 0.00606 -0.0663 0.7455 1.0000
3.500 0.6631 0.01298 0.00623 -0.0646 0.7084 1.0000
3.750 0.6868 0.01321 0.00641 -0.0628 0.6684 1.0000
4.000 0.7098 0.01350 0.00660 -0.0610 0.6233 1.0000
4.250 0.7315 0.01383 0.00680 -0.0589 0.5703 1.0000
4.500 0.7515 0.01420 0.00692 -0.0566 0.5057 1.0000
4.750 0.7708 0.01462 0.00707 -0.0546 0.4209 1.0000
5.000 0.7878 0.01565 0.00754 -0.0525 0.2975 1.0000
5.250 0.7966 0.01915 0.00944 -0.0498 0.0813 1.0000
5.500 0.8164 0.02075 0.01098 -0.0483 0.0669 1.0000
5.750 0.8371 0.02237 0.01270 -0.0467 0.0612 1.0000
6.000 0.8596 0.02411 0.01442 -0.0454 0.0570 1.0000
6.250 0.8830 0.02668 0.01689 -0.0445 0.0524 1.0000
6.500 0.9090 0.02857 0.01904 -0.0435 0.0501 1.0000
6.750 0.9347 0.03113 0.02188 -0.0425 0.0496 1.0000
7.000 0.9587 0.03412 0.02525 -0.0412 0.0499 1.0000
7.250 0.9800 0.03768 0.02924 -0.0399 0.0510 1.0000
7.500 0.9984 0.04221 0.03417 -0.0387 0.0525 1.0000
7.750 1.0160 0.04659 0.03890 -0.0376 0.0534 1.0000
8.250 1.0344 0.05558 0.04958 -0.0326 0.0684 1.0000
10.000 0.7487 0.11429 0.11010 -0.0552 0.1517 1.0000
10.250 0.7475 0.11908 0.11488 -0.0564 0.1484 1.0000
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