USA 40 B AIRFOIL (usa40b-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 40 B AIRFOIL (usa40b-il) Reynolds number: 100,000 Max Cl/Cd: 39.93 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa40b-il-100000.txt Download as CSV file: xf-usa40b-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.2610 0.09768 0.09278 -0.0272 1.0000 0.1392
-8.000 -0.2995 0.09728 0.09256 -0.0292 1.0000 0.1441
-7.750 -0.3575 0.09743 0.09294 -0.0297 1.0000 0.1448
-7.500 -0.2822 0.09107 0.08647 -0.0253 1.0000 0.1484
-7.250 -0.2785 0.08931 0.08479 -0.0230 1.0000 0.1520
-7.000 -0.2945 0.08821 0.08381 -0.0206 1.0000 0.1561
-6.750 -0.3329 0.08813 0.08391 -0.0171 1.0000 0.1583
-6.500 -0.4110 0.08676 0.08265 -0.0265 1.0000 0.1606
-6.250 -0.3927 0.08458 0.08061 -0.0168 1.0000 0.1620
-6.000 -0.3700 0.08296 0.07903 -0.0121 0.9981 0.1649
-5.750 -0.3343 0.07575 0.07175 -0.0309 0.9839 0.1790
-5.500 -0.2976 0.07322 0.06924 -0.0298 0.9765 0.1830
-5.250 -0.2664 0.06701 0.06292 -0.0438 0.9631 0.1975
-5.000 -0.2300 0.03890 0.03251 -0.0755 0.9526 0.1025
-4.750 -0.1885 0.03530 0.02852 -0.0796 0.9431 0.1020
-4.500 -0.1492 0.03233 0.02515 -0.0826 0.9328 0.1016
-4.250 -0.1006 0.02949 0.02186 -0.0867 0.9254 0.1014
-4.000 -0.0617 0.02753 0.01950 -0.0885 0.9137 0.1022
-3.750 -0.0090 0.02535 0.01716 -0.0927 0.9076 0.1060
-3.500 0.0300 0.02409 0.01584 -0.0942 0.8949 0.1110
-3.250 0.0783 0.02257 0.01408 -0.0969 0.8857 0.1164
-3.000 0.1199 0.02115 0.01281 -0.0985 0.8735 0.1250
-2.750 0.1562 0.02007 0.01181 -0.0992 0.8583 0.1373
-2.500 0.1922 0.01907 0.01089 -0.0996 0.8423 0.1562
-2.250 0.2263 0.01817 0.01014 -0.0997 0.8243 0.1889
-2.000 0.2574 0.01704 0.00934 -0.0994 0.8046 0.2553
-1.750 0.2865 0.01608 0.00885 -0.0989 0.7832 0.3664
-1.500 0.3147 0.01543 0.00854 -0.0978 0.7601 0.4685
-1.250 0.3366 0.01514 0.00857 -0.0954 0.7327 0.5708
-1.000 0.3591 0.01514 0.00863 -0.0926 0.7056 0.6592
-0.750 0.3781 0.01528 0.00872 -0.0893 0.6747 0.7214
-0.500 0.3960 0.01537 0.00868 -0.0857 0.6451 0.7755
-0.250 0.4109 0.01531 0.00854 -0.0816 0.6142 0.8263
0.000 0.4273 0.01511 0.00821 -0.0779 0.5855 0.8764
0.250 0.4639 0.01495 0.00782 -0.0783 0.5549 0.9340
0.500 0.5151 0.01516 0.00768 -0.0829 0.5225 1.0000
0.750 0.5395 0.01552 0.00770 -0.0826 0.5023 1.0000
1.000 0.5660 0.01593 0.00781 -0.0827 0.4842 1.0000
1.250 0.5931 0.01636 0.00799 -0.0827 0.4680 1.0000
1.500 0.6205 0.01682 0.00821 -0.0828 0.4536 1.0000
1.750 0.6484 0.01731 0.00845 -0.0828 0.4413 1.0000
2.000 0.6760 0.01778 0.00872 -0.0828 0.4296 1.0000
2.250 0.7034 0.01830 0.00912 -0.0828 0.4193 1.0000
2.500 0.7316 0.01884 0.00943 -0.0829 0.4104 1.0000
2.750 0.7584 0.01939 0.00992 -0.0827 0.4016 1.0000
3.000 0.7860 0.01992 0.01029 -0.0826 0.3935 1.0000
3.250 0.8131 0.02054 0.01081 -0.0825 0.3862 1.0000
3.500 0.8393 0.02108 0.01132 -0.0822 0.3789 1.0000
3.750 0.8677 0.02173 0.01178 -0.0823 0.3733 1.0000
4.000 0.8934 0.02242 0.01250 -0.0820 0.3677 1.0000
4.250 0.9188 0.02306 0.01316 -0.0816 0.3619 1.0000
4.500 0.9459 0.02370 0.01370 -0.0815 0.3569 1.0000
4.750 0.9730 0.02452 0.01443 -0.0814 0.3524 1.0000
5.000 0.9960 0.02524 0.01528 -0.0807 0.3472 1.0000
5.250 1.0208 0.02596 0.01605 -0.0803 0.3427 1.0000
5.500 1.0470 0.02672 0.01676 -0.0801 0.3389 1.0000
5.750 1.0748 0.02772 0.01763 -0.0802 0.3357 1.0000
6.000 1.0951 0.02871 0.01885 -0.0792 0.3323 1.0000
6.250 1.1163 0.02974 0.02004 -0.0784 0.3288 1.0000
6.500 1.1392 0.03070 0.02107 -0.0778 0.3253 1.0000
6.750 1.1633 0.03159 0.02198 -0.0774 0.3220 1.0000
7.000 1.1890 0.03256 0.02290 -0.0772 0.3191 1.0000
7.250 1.2096 0.03396 0.02440 -0.0765 0.3165 1.0000
7.500 1.2237 0.03551 0.02625 -0.0750 0.3139 1.0000
7.750 1.2380 0.03716 0.02813 -0.0736 0.3113 1.0000
8.000 1.2529 0.03883 0.02997 -0.0724 0.3088 1.0000
8.250 1.2688 0.04045 0.03174 -0.0713 0.3066 1.0000
8.500 1.2888 0.04178 0.03314 -0.0706 0.3043 1.0000
8.750 1.3126 0.04294 0.03429 -0.0703 0.3020 1.0000
9.000 1.3344 0.04472 0.03605 -0.0700 0.2998 1.0000
9.250 1.3230 0.04807 0.03984 -0.0666 0.2980 1.0000
9.500 1.2988 0.05255 0.04471 -0.0628 0.2964 1.0000
9.750 1.0254 0.08332 0.07613 -0.0597 0.2995 1.0000
10.000 0.8120 0.12522 0.11829 -0.0832 0.3549 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 40 B AIRFOIL (usa40b-il)