USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 1,000,000 Max Cl/Cd: 129.19 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa41-il-1000000.txt Download as CSV file: xf-usa41-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: USA 41 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3487 0.08614 0.08460 -0.0235 1.0000 0.0122
-7.500 -0.3539 0.08404 0.08254 -0.0222 1.0000 0.0122
-7.250 -0.3551 0.08206 0.08060 -0.0218 0.9994 0.0130
-7.000 -0.3210 0.07486 0.07338 -0.0357 0.9950 0.0147
-6.750 -0.2948 0.06923 0.06774 -0.0442 0.9903 0.0148
-6.500 -0.2650 0.06343 0.06190 -0.0533 0.9857 0.0148
-6.250 -0.2347 0.05678 0.05520 -0.0631 0.9797 0.0148
-6.000 -0.1854 0.01919 0.01619 -0.0994 0.9658 0.0151
-5.750 -0.1542 0.01818 0.01507 -0.1010 0.9603 0.0157
-5.500 -0.1208 0.01842 0.01535 -0.1023 0.9554 0.0162
-5.250 -0.0903 0.01789 0.01473 -0.1032 0.9470 0.0171
-5.000 -0.0606 0.01630 0.01281 -0.1040 0.9378 0.0186
-4.500 -0.0076 0.01332 0.00922 -0.1041 0.9171 0.0210
-4.250 0.0189 0.01302 0.00889 -0.1040 0.9071 0.0219
-4.000 0.0454 0.01271 0.00849 -0.1037 0.8975 0.0232
-3.750 0.0714 0.01215 0.00777 -0.1032 0.8873 0.0244
-3.500 0.0975 0.01151 0.00696 -0.1026 0.8775 0.0251
-3.250 0.1241 0.01120 0.00651 -0.1022 0.8677 0.0259
-3.000 0.1491 0.01001 0.00508 -0.1016 0.8570 0.0278
-2.750 0.1748 0.00935 0.00436 -0.1011 0.8455 0.0288
-2.500 0.2009 0.00887 0.00378 -0.1005 0.8329 0.0292
-2.250 0.2269 0.00844 0.00326 -0.1000 0.8187 0.0294
-2.000 0.2528 0.00808 0.00281 -0.0993 0.8021 0.0297
-1.500 0.3043 0.00761 0.00214 -0.0981 0.7634 0.0314
-1.250 0.3298 0.00745 0.00187 -0.0974 0.7416 0.0322
-1.000 0.3555 0.00733 0.00165 -0.0968 0.7196 0.0329
-0.750 0.3812 0.00727 0.00147 -0.0962 0.6974 0.0334
-0.500 0.4071 0.00722 0.00132 -0.0956 0.6741 0.0338
-0.250 0.4329 0.00721 0.00120 -0.0950 0.6500 0.0341
0.000 0.4588 0.00722 0.00110 -0.0945 0.6264 0.0345
0.250 0.4848 0.00726 0.00104 -0.0940 0.6030 0.0352
0.750 0.5370 0.00734 0.00092 -0.0930 0.5628 0.0381
1.000 0.5634 0.00738 0.00091 -0.0926 0.5459 0.0450
1.250 0.5890 0.00722 0.00099 -0.0922 0.5310 0.1575
1.500 0.6155 0.00729 0.00107 -0.0918 0.5175 0.1836
1.750 0.6420 0.00737 0.00114 -0.0915 0.5043 0.1976
2.000 0.6684 0.00745 0.00120 -0.0911 0.4907 0.2106
2.250 0.6948 0.00752 0.00126 -0.0908 0.4770 0.2221
2.750 0.7468 0.00751 0.00142 -0.0900 0.4502 0.3231
3.000 0.7900 0.00633 0.00167 -0.0940 0.4352 1.0000
3.250 0.8163 0.00646 0.00175 -0.0936 0.4231 1.0000
3.500 0.8424 0.00659 0.00186 -0.0932 0.4120 1.0000
3.750 0.8684 0.00674 0.00196 -0.0928 0.3998 1.0000
4.000 0.8940 0.00692 0.00208 -0.0923 0.3830 1.0000
4.250 0.9189 0.00716 0.00221 -0.0917 0.3598 1.0000
4.500 0.9435 0.00744 0.00238 -0.0911 0.3336 1.0000
4.750 0.9688 0.00765 0.00254 -0.0906 0.3173 1.0000
5.000 0.9939 0.00788 0.00271 -0.0901 0.2985 1.0000
5.250 1.0181 0.00820 0.00292 -0.0894 0.2717 1.0000
5.500 1.0408 0.00867 0.00320 -0.0885 0.2301 1.0000
5.750 1.0603 0.00950 0.00366 -0.0872 0.1626 1.0000
6.000 1.0749 0.01091 0.00450 -0.0852 0.0612 1.0000
6.250 1.0944 0.01180 0.00516 -0.0838 0.0227 1.0000
6.500 1.1180 0.01219 0.00558 -0.0830 0.0187 1.0000
6.750 1.1409 0.01265 0.00606 -0.0821 0.0162 1.0000
7.000 1.1629 0.01323 0.00671 -0.0810 0.0144 1.0000
7.250 1.1860 0.01363 0.00715 -0.0802 0.0135 1.0000
7.500 1.2084 0.01409 0.00765 -0.0793 0.0126 1.0000
7.750 1.2302 0.01460 0.00820 -0.0783 0.0116 1.0000
8.000 1.2462 0.01571 0.00943 -0.0763 0.0104 1.0000
8.250 1.2669 0.01628 0.01006 -0.0751 0.0100 1.0000
8.500 1.2870 0.01687 0.01071 -0.0739 0.0096 1.0000
8.750 1.3060 0.01754 0.01145 -0.0725 0.0092 1.0000
9.000 1.3243 0.01823 0.01220 -0.0710 0.0088 1.0000
9.250 1.3411 0.01903 0.01308 -0.0693 0.0085 1.0000
9.500 1.3586 0.01971 0.01380 -0.0678 0.0080 1.0000
9.750 1.3740 0.02053 0.01466 -0.0660 0.0077 1.0000
10.000 1.3783 0.02224 0.01648 -0.0626 0.0072 1.0000
10.250 1.3843 0.02365 0.01802 -0.0593 0.0069 1.0000
10.500 1.3953 0.02442 0.01886 -0.0568 0.0068 1.0000
10.750 1.4034 0.02544 0.01999 -0.0540 0.0066 1.0000
11.000 1.4112 0.02651 0.02115 -0.0513 0.0064 1.0000
11.250 1.4168 0.02784 0.02259 -0.0485 0.0063 1.0000
11.500 1.4206 0.02943 0.02430 -0.0458 0.0061 1.0000
11.750 1.4247 0.03106 0.02605 -0.0435 0.0061 1.0000
12.000 1.4297 0.03253 0.02762 -0.0414 0.0058 1.0000
12.250 1.4338 0.03417 0.02936 -0.0396 0.0057 1.0000
12.500 1.4337 0.03646 0.03179 -0.0376 0.0056 1.0000
12.750 1.4355 0.03850 0.03395 -0.0362 0.0056 1.0000
13.000 1.4394 0.04021 0.03574 -0.0353 0.0054 1.0000
13.250 1.4319 0.04378 0.03953 -0.0338 0.0054 1.0000
13.500 1.4320 0.04621 0.04204 -0.0334 0.0053 1.0000
13.750 1.4332 0.04863 0.04454 -0.0334 0.0051 1.0000
14.000 1.4258 0.05247 0.04854 -0.0333 0.0051 1.0000
14.250 1.4192 0.05634 0.05254 -0.0338 0.0051 1.0000
14.500 1.4133 0.06021 0.05650 -0.0347 0.0049 1.0000
14.750 1.4021 0.06516 0.06163 -0.0359 0.0050 1.0000
15.000 1.3892 0.07061 0.06724 -0.0377 0.0049 1.0000
15.250 1.3731 0.07683 0.07362 -0.0400 0.0049 1.0000
15.500 1.3570 0.08336 0.08031 -0.0428 0.0048 1.0000
15.750 1.3390 0.09058 0.08770 -0.0462 0.0048 1.0000
16.000 1.3270 0.09702 0.09428 -0.0494 0.0049 1.0000
16.250 1.3090 0.10490 0.10232 -0.0536 0.0049 1.0000
16.500 1.2901 0.11337 0.11094 -0.0583 0.0049 1.0000
16.750 1.2691 0.12265 0.12037 -0.0636 0.0048 1.0000
17.000 1.2571 0.13037 0.12823 -0.0682 0.0049 1.0000
17.250 1.2364 0.14060 0.13862 -0.0743 0.0050 1.0000
17.500 1.2176 0.15076 0.14890 -0.0806 0.0049 1.0000
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Polar data table (+)
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