USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 1,000,000 Max Cl/Cd: 122.82 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa41-il-1000000-n5.txt Download as CSV file: xf-usa41-il-1000000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3666 0.09841 0.09678 -0.0238 1.0000 0.0042
-9.000 -0.3697 0.09372 0.09212 -0.0250 1.0000 0.0046
-8.000 -0.3888 0.01635 0.01308 -0.1040 0.9624 0.0061
-7.750 -0.3574 0.01438 0.01077 -0.1059 0.9537 0.0065
-7.500 -0.3253 0.01312 0.00924 -0.1074 0.9425 0.0068
-7.250 -0.2945 0.01229 0.00823 -0.1083 0.9300 0.0075
-7.000 -0.2668 0.01157 0.00735 -0.1085 0.9193 0.0083
-6.500 -0.2138 0.01042 0.00588 -0.1082 0.9018 0.0105
-6.250 -0.1862 0.01039 0.00585 -0.1080 0.8940 0.0120
-6.000 -0.1592 0.01026 0.00566 -0.1077 0.8861 0.0133
-5.750 -0.1323 0.01010 0.00543 -0.1075 0.8784 0.0144
-5.500 -0.1045 0.01032 0.00566 -0.1072 0.8699 0.0152
-5.250 -0.0769 0.01049 0.00581 -0.1070 0.8615 0.0162
-5.000 -0.0499 0.01048 0.00572 -0.1067 0.8523 0.0175
-4.750 -0.0231 0.01034 0.00549 -0.1064 0.8417 0.0187
-4.500 0.0038 0.01027 0.00533 -0.1060 0.8302 0.0194
-4.250 0.0307 0.01037 0.00535 -0.1057 0.8162 0.0199
-4.000 0.0572 0.01036 0.00524 -0.1053 0.7995 0.0201
-3.750 0.0815 0.00947 0.00418 -0.1047 0.7805 0.0216
-3.500 0.1074 0.00930 0.00391 -0.1042 0.7603 0.0223
-3.250 0.1332 0.00915 0.00365 -0.1037 0.7394 0.0229
-3.000 0.1592 0.00902 0.00340 -0.1032 0.7193 0.0237
-2.750 0.1850 0.00881 0.00307 -0.1027 0.6989 0.0240
-2.250 0.2369 0.00847 0.00249 -0.1017 0.6578 0.0246
-2.000 0.2627 0.00833 0.00223 -0.1012 0.6351 0.0247
-1.750 0.2889 0.00823 0.00202 -0.1007 0.6138 0.0250
-1.500 0.3148 0.00817 0.00184 -0.1002 0.5912 0.0254
-1.250 0.3410 0.00808 0.00164 -0.0998 0.5718 0.0254
-1.000 0.3674 0.00801 0.00148 -0.0994 0.5548 0.0255
-0.750 0.3939 0.00796 0.00134 -0.0990 0.5395 0.0257
-0.500 0.4204 0.00792 0.00123 -0.0987 0.5259 0.0261
-0.250 0.4471 0.00789 0.00115 -0.0983 0.5142 0.0267
0.250 0.5007 0.00790 0.00106 -0.0977 0.4914 0.0281
0.500 0.5276 0.00792 0.00105 -0.0975 0.4816 0.0286
0.750 0.5543 0.00793 0.00102 -0.0972 0.4718 0.0299
1.000 0.5812 0.00795 0.00100 -0.0969 0.4612 0.0310
1.250 0.6080 0.00799 0.00100 -0.0966 0.4514 0.0323
1.500 0.6347 0.00804 0.00102 -0.0963 0.4414 0.0337
1.750 0.6614 0.00810 0.00106 -0.0960 0.4306 0.0354
2.000 0.6880 0.00815 0.00111 -0.0957 0.4190 0.0387
2.250 0.7138 0.00803 0.00123 -0.0954 0.4058 0.1648
2.500 0.7400 0.00814 0.00133 -0.0950 0.3900 0.1804
2.750 0.7662 0.00826 0.00142 -0.0947 0.3757 0.1904
3.000 0.7925 0.00836 0.00152 -0.0944 0.3641 0.1986
3.500 0.8452 0.00855 0.00173 -0.0937 0.3450 0.2141
3.750 0.8714 0.00863 0.00185 -0.0934 0.3350 0.2355
4.000 0.8926 0.00800 0.00214 -0.0925 0.3241 0.6844
4.250 0.9359 0.00762 0.00234 -0.0961 0.3097 1.0000
4.500 0.9598 0.00797 0.00253 -0.0954 0.2762 1.0000
4.750 0.9840 0.00830 0.00273 -0.0948 0.2492 1.0000
5.000 1.0069 0.00876 0.00301 -0.0940 0.2111 1.0000
5.250 1.0285 0.00937 0.00338 -0.0930 0.1629 1.0000
5.500 1.0466 0.01037 0.00397 -0.0915 0.0874 1.0000
6.000 1.0895 0.01165 0.00492 -0.0894 0.0215 1.0000
6.250 1.1131 0.01204 0.00527 -0.0886 0.0147 1.0000
6.500 1.1371 0.01235 0.00561 -0.0880 0.0118 1.0000
6.750 1.1608 0.01270 0.00597 -0.0873 0.0102 1.0000
7.000 1.1836 0.01314 0.00643 -0.0864 0.0083 1.0000
7.250 1.2071 0.01348 0.00680 -0.0857 0.0076 1.0000
7.500 1.2301 0.01385 0.00720 -0.0849 0.0069 1.0000
7.750 1.2523 0.01429 0.00767 -0.0840 0.0061 1.0000
8.000 1.2734 0.01485 0.00827 -0.0829 0.0055 1.0000
8.250 1.2955 0.01528 0.00875 -0.0820 0.0052 1.0000
8.500 1.3168 0.01576 0.00929 -0.0810 0.0049 1.0000
8.750 1.3375 0.01628 0.00987 -0.0799 0.0046 1.0000
9.000 1.3579 0.01680 0.01043 -0.0788 0.0043 1.0000
9.250 1.3778 0.01733 0.01099 -0.0776 0.0040 1.0000
9.500 1.3951 0.01808 0.01180 -0.0760 0.0037 1.0000
9.750 1.4124 0.01877 0.01257 -0.0745 0.0035 1.0000
10.000 1.4295 0.01944 0.01332 -0.0729 0.0034 1.0000
10.250 1.4456 0.02014 0.01409 -0.0712 0.0032 1.0000
10.500 1.4601 0.02091 0.01497 -0.0693 0.0031 1.0000
10.750 1.4727 0.02168 0.01581 -0.0670 0.0030 1.0000
11.000 1.4811 0.02253 0.01675 -0.0641 0.0029 1.0000
11.250 1.4907 0.02333 0.01762 -0.0615 0.0028 1.0000
11.500 1.4988 0.02423 0.01860 -0.0588 0.0027 1.0000
11.750 1.5039 0.02538 0.01985 -0.0559 0.0026 1.0000
12.000 1.5126 0.02634 0.02088 -0.0538 0.0025 1.0000
12.250 1.5166 0.02768 0.02232 -0.0513 0.0025 1.0000
12.500 1.5215 0.02905 0.02377 -0.0492 0.0024 1.0000
12.750 1.5226 0.03082 0.02565 -0.0470 0.0024 1.0000
13.000 1.5238 0.03270 0.02765 -0.0452 0.0023 1.0000
13.250 1.5199 0.03521 0.03029 -0.0434 0.0022 1.0000
13.500 1.5101 0.03852 0.03376 -0.0419 0.0022 1.0000
13.750 1.5090 0.04117 0.03653 -0.0411 0.0021 1.0000
14.000 1.5080 0.04394 0.03944 -0.0407 0.0021 1.0000
14.250 1.5034 0.04735 0.04297 -0.0406 0.0021 1.0000
14.500 1.4984 0.05097 0.04673 -0.0408 0.0021 1.0000
14.750 1.4966 0.05432 0.05020 -0.0413 0.0021 1.0000
15.000 1.4880 0.05880 0.05483 -0.0423 0.0020 1.0000
15.250 1.4791 0.06347 0.05963 -0.0435 0.0020 1.0000
15.500 1.4708 0.06826 0.06456 -0.0450 0.0020 1.0000
15.750 1.4598 0.07365 0.07009 -0.0468 0.0020 1.0000
16.000 1.4465 0.07955 0.07613 -0.0490 0.0020 1.0000
16.250 1.4383 0.08482 0.08153 -0.0511 0.0020 1.0000
16.500 1.4260 0.09097 0.08781 -0.0537 0.0019 1.0000
16.750 1.4141 0.09719 0.09415 -0.0565 0.0019 1.0000
17.000 1.4021 0.10367 0.10075 -0.0595 0.0019 1.0000
17.250 1.3833 0.11166 0.10890 -0.0634 0.0019 1.0000
17.500 1.3697 0.11883 0.11619 -0.0670 0.0019 1.0000
17.750 1.3536 0.12678 0.12429 -0.0712 0.0019 1.0000
18.000 1.3372 0.13503 0.13267 -0.0757 0.0019 1.0000
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