USA 40 B AIRFOIL (usa40b-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 40 B AIRFOIL (usa40b-il) Reynolds number: 50,000 Max Cl/Cd: 25.06 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40b-il-50000.txt Download as CSV file: xf-usa40b-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: USA 40 B 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2811 0.10872 0.10205 -0.0191 1.0000 0.2697
-7.750 -0.2365 0.10155 0.09485 -0.0176 1.0000 0.2773
-7.500 -0.2541 0.10148 0.09493 -0.0164 1.0000 0.2872
-7.250 -0.2423 0.09776 0.09130 -0.0151 1.0000 0.2934
-7.000 -0.2459 0.09641 0.09007 -0.0132 1.0000 0.3043
-6.750 -0.2582 0.09485 0.08867 -0.0111 1.0000 0.3104
-6.500 -0.2476 0.09229 0.08619 -0.0090 1.0000 0.3205
-6.250 -0.3067 0.09511 0.08930 -0.0036 1.0000 0.3262
-6.000 -0.2641 0.08933 0.08351 -0.0033 1.0000 0.3356
-5.750 -0.3066 0.09076 0.08516 0.0022 1.0000 0.3439
-5.500 -0.3107 0.08850 0.08303 0.0051 1.0000 0.3497
-5.250 -0.3160 0.08720 0.08182 0.0084 1.0000 0.3588
-5.000 -0.3492 0.08711 0.08191 0.0116 1.0000 0.3673
-4.750 -0.3422 0.08493 0.07979 0.0147 1.0000 0.3767
-4.500 -0.3633 0.08400 0.07900 0.0171 1.0000 0.3880
-4.250 -0.3863 0.08377 0.07889 0.0183 1.0000 0.4044
-4.000 -0.3721 0.08084 0.07602 0.0224 1.0000 0.4117
-3.750 -0.3235 0.05288 0.04647 -0.0383 1.0000 0.1803
-3.500 -0.2883 0.04853 0.04166 -0.0443 0.9966 0.1734
-3.250 -0.2204 0.04388 0.03636 -0.0550 0.9809 0.1698
-3.000 -0.1585 0.04036 0.03220 -0.0631 0.9632 0.1695
-2.750 -0.0964 0.03797 0.02906 -0.0701 0.9445 0.1765
-2.500 -0.0364 0.03603 0.02705 -0.0759 0.9253 0.1862
-2.250 0.0193 0.03439 0.02520 -0.0804 0.9042 0.2027
-2.000 0.0716 0.03287 0.02358 -0.0839 0.8816 0.2279
-1.750 0.1262 0.03114 0.02203 -0.0873 0.8598 0.2738
-1.500 0.1841 0.02909 0.02055 -0.0904 0.8394 0.3823
-1.250 0.2292 0.02728 0.01982 -0.0899 0.8191 0.5622
-1.000 0.2592 0.02653 0.01954 -0.0853 0.7986 0.6999
-0.750 0.2728 0.02586 0.01902 -0.0784 0.7756 0.8064
-0.500 0.4068 0.02315 0.01601 -0.0903 0.7437 1.0000
-0.250 0.4186 0.02319 0.01573 -0.0885 0.7155 1.0000
0.000 0.4466 0.02320 0.01534 -0.0883 0.6889 1.0000
0.250 0.4759 0.02338 0.01513 -0.0881 0.6639 1.0000
0.500 0.5002 0.02383 0.01530 -0.0875 0.6390 1.0000
0.750 0.5275 0.02422 0.01539 -0.0872 0.6174 1.0000
1.000 0.5549 0.02466 0.01554 -0.0869 0.5979 1.0000
1.250 0.5819 0.02518 0.01581 -0.0865 0.5803 1.0000
1.500 0.6086 0.02576 0.01616 -0.0862 0.5643 1.0000
1.750 0.6363 0.02633 0.01649 -0.0859 0.5500 1.0000
2.000 0.6652 0.02685 0.01675 -0.0858 0.5370 1.0000
2.250 0.6872 0.02783 0.01767 -0.0852 0.5238 1.0000
2.500 0.7128 0.02870 0.01839 -0.0848 0.5130 1.0000
2.750 0.7390 0.02949 0.01902 -0.0845 0.5025 1.0000
3.000 0.7601 0.03069 0.02020 -0.0839 0.4923 1.0000
3.250 0.7874 0.03149 0.02084 -0.0837 0.4837 1.0000
3.500 0.8056 0.03304 0.02244 -0.0830 0.4757 1.0000
3.750 0.8281 0.03427 0.02363 -0.0825 0.4682 1.0000
4.000 0.8537 0.03539 0.02463 -0.0822 0.4618 1.0000
4.250 0.8627 0.03763 0.02707 -0.0810 0.4543 1.0000
4.500 0.8874 0.03877 0.02812 -0.0806 0.4482 1.0000
4.750 0.9073 0.04048 0.02982 -0.0801 0.4433 1.0000
5.000 0.9016 0.04422 0.03384 -0.0785 0.4385 1.0000
5.250 0.8996 0.04776 0.03752 -0.0773 0.4342 1.0000
5.500 0.9132 0.05003 0.03979 -0.0766 0.4300 1.0000
5.750 0.9426 0.05116 0.04084 -0.0765 0.4261 1.0000
6.000 0.8712 0.06113 0.05112 -0.0745 0.4249 1.0000
6.250 0.8208 0.07012 0.06019 -0.0748 0.4269 1.0000
6.500 0.5459 0.10135 0.09187 -0.0925 0.5976 1.0000
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Polar data table (+)
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