USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 500,000 Max Cl/Cd: 107.8 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa41-il-500000.txt Download as CSV file: xf-usa41-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3467 0.08894 0.08678 -0.0261 1.0000 0.0195
-7.500 -0.3531 0.08683 0.08473 -0.0251 1.0000 0.0195
-7.250 -0.3553 0.08427 0.08222 -0.0254 1.0000 0.0196
-7.000 -0.3610 0.08212 0.08011 -0.0246 1.0000 0.0196
-6.750 -0.3424 0.07742 0.07540 -0.0308 0.9978 0.0197
-6.500 -0.3209 0.07015 0.06812 -0.0389 0.9944 0.0202
-6.250 -0.2988 0.06725 0.06520 -0.0416 0.9909 0.0208
-6.000 -0.2697 0.06353 0.06145 -0.0474 0.9872 0.0217
-5.750 -0.2339 0.05866 0.05653 -0.0562 0.9841 0.0228
-5.500 -0.1822 0.05178 0.04951 -0.0701 0.9779 0.0262
-5.250 -0.1392 0.04497 0.04251 -0.0797 0.9738 0.0264
-4.750 -0.0764 0.03428 0.03153 -0.0911 0.9649 0.0291
-4.500 -0.0397 0.03063 0.02767 -0.0953 0.9607 0.0321
-4.250 0.0004 0.02218 0.01840 -0.1011 0.9564 0.0375
-4.000 0.0279 0.02114 0.01739 -0.1017 0.9485 0.0395
-3.750 0.0655 0.02199 0.01796 -0.1028 0.9432 0.0470
-3.500 0.0892 0.01801 0.01382 -0.1040 0.9335 0.0522
-3.250 0.1188 0.01686 0.01240 -0.1047 0.9250 0.0628
-3.000 0.1489 0.01297 0.00777 -0.1038 0.9156 0.0426
-2.750 0.1761 0.01152 0.00598 -0.1031 0.9050 0.0397
-2.500 0.2034 0.01075 0.00504 -0.1026 0.8945 0.0391
-2.250 0.2304 0.01016 0.00432 -0.1021 0.8839 0.0393
-2.000 0.2565 0.00970 0.00378 -0.1014 0.8718 0.0404
-1.750 0.2824 0.00926 0.00326 -0.1007 0.8592 0.0407
-1.500 0.3084 0.00890 0.00284 -0.1000 0.8458 0.0412
-1.250 0.3344 0.00862 0.00249 -0.0993 0.8313 0.0420
-1.000 0.3604 0.00840 0.00219 -0.0987 0.8151 0.0427
-0.750 0.3864 0.00824 0.00194 -0.0980 0.7975 0.0435
-0.500 0.4122 0.00813 0.00175 -0.0973 0.7770 0.0446
-0.250 0.4379 0.00808 0.00158 -0.0966 0.7554 0.0461
0.000 0.4637 0.00806 0.00144 -0.0959 0.7323 0.0473
0.250 0.4892 0.00807 0.00134 -0.0951 0.7090 0.0531
0.500 0.5140 0.00789 0.00137 -0.0944 0.6847 0.1480
0.750 0.5395 0.00799 0.00143 -0.0938 0.6606 0.1841
1.000 0.5649 0.00810 0.00147 -0.0932 0.6370 0.2009
1.250 0.5904 0.00820 0.00151 -0.0926 0.6138 0.2162
1.500 0.6157 0.00828 0.00155 -0.0920 0.5931 0.2340
1.750 0.6411 0.00830 0.00160 -0.0915 0.5740 0.2622
2.000 0.6593 0.00738 0.00182 -0.0898 0.5579 0.7193
2.250 0.7083 0.00710 0.00189 -0.0945 0.5393 1.0000
2.500 0.7339 0.00728 0.00199 -0.0939 0.5241 1.0000
2.750 0.7596 0.00746 0.00208 -0.0934 0.5095 1.0000
3.000 0.7853 0.00763 0.00219 -0.0929 0.4953 1.0000
3.250 0.8110 0.00780 0.00231 -0.0924 0.4813 1.0000
3.500 0.8366 0.00798 0.00245 -0.0919 0.4673 1.0000
3.750 0.8623 0.00815 0.00258 -0.0914 0.4539 1.0000
4.000 0.8880 0.00833 0.00273 -0.0909 0.4414 1.0000
4.250 0.9136 0.00851 0.00289 -0.0904 0.4296 1.0000
4.500 0.9389 0.00871 0.00307 -0.0898 0.4143 1.0000
4.750 0.9632 0.00897 0.00322 -0.0891 0.3893 1.0000
5.000 0.9877 0.00922 0.00341 -0.0885 0.3655 1.0000
5.250 1.0122 0.00950 0.00362 -0.0879 0.3446 1.0000
5.500 1.0366 0.00978 0.00387 -0.0872 0.3237 1.0000
5.750 1.0600 0.01016 0.00413 -0.0864 0.2933 1.0000
6.000 1.0815 0.01075 0.00448 -0.0854 0.2414 1.0000
6.250 1.0949 0.01228 0.00531 -0.0834 0.1208 1.0000
6.500 1.1064 0.01414 0.00658 -0.0809 0.0307 1.0000
6.750 1.1283 0.01474 0.00727 -0.0798 0.0257 1.0000
7.000 1.1482 0.01555 0.00813 -0.0784 0.0218 1.0000
7.250 1.1675 0.01639 0.00908 -0.0769 0.0200 1.0000
7.500 1.1876 0.01708 0.00986 -0.0756 0.0188 1.0000
7.750 1.2064 0.01788 0.01073 -0.0741 0.0176 1.0000
8.000 1.2240 0.01875 0.01168 -0.0725 0.0166 1.0000
8.250 1.2375 0.01998 0.01297 -0.0703 0.0154 1.0000
8.500 1.2451 0.02184 0.01495 -0.0672 0.0145 1.0000
8.750 1.2624 0.02266 0.01586 -0.0656 0.0141 1.0000
9.000 1.2774 0.02374 0.01703 -0.0636 0.0136 1.0000
9.250 1.2910 0.02503 0.01843 -0.0615 0.0132 1.0000
9.500 1.3048 0.02633 0.01984 -0.0595 0.0127 1.0000
9.750 1.3179 0.02779 0.02140 -0.0575 0.0124 1.0000
10.000 1.3306 0.02933 0.02304 -0.0554 0.0120 1.0000
10.250 1.3424 0.03090 0.02471 -0.0533 0.0117 1.0000
10.500 1.3519 0.03223 0.02613 -0.0510 0.0112 1.0000
10.750 1.3618 0.03413 0.02812 -0.0490 0.0109 1.0000
11.000 1.3747 0.03916 0.03339 -0.0484 0.0103 1.0000
11.250 1.3784 0.04253 0.03702 -0.0459 0.0103 1.0000
11.500 1.3787 0.04481 0.03950 -0.0430 0.0103 1.0000
11.750 1.3749 0.04757 0.04247 -0.0401 0.0103 1.0000
12.000 1.3659 0.05092 0.04606 -0.0373 0.0103 1.0000
12.250 1.3560 0.05407 0.04943 -0.0349 0.0103 1.0000
12.500 1.3503 0.05637 0.05187 -0.0332 0.0104 1.0000
12.750 1.3391 0.05979 0.05549 -0.0320 0.0104 1.0000
13.000 1.3227 0.06425 0.06016 -0.0315 0.0104 1.0000
13.250 1.3191 0.06649 0.06254 -0.0314 0.0106 1.0000
13.500 1.3022 0.07135 0.06762 -0.0323 0.0107 1.0000
14.000 1.0819 0.09687 0.09421 -0.0418 0.0125 1.0000
14.250 1.0524 0.10450 0.10201 -0.0466 0.0125 1.0000
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