USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 5 AIRFOIL (usa5-il) Reynolds number: 200,000 Max Cl/Cd: 70.96 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa5-il-200000-n5.txt Download as CSV file: xf-usa5-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 5 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3260 0.09861 0.09539 -0.0260 1.0000 0.0185
-7.500 -0.3323 0.09681 0.09366 -0.0244 1.0000 0.0185
-7.250 -0.3380 0.09480 0.09172 -0.0231 1.0000 0.0185
-7.000 -0.3406 0.09253 0.08949 -0.0223 1.0000 0.0185
-6.750 -0.3315 0.08913 0.08611 -0.0248 0.9981 0.0186
-6.500 -0.3085 0.08454 0.08150 -0.0307 0.9936 0.0186
-6.250 -0.2845 0.07979 0.07673 -0.0366 0.9888 0.0186
-6.000 -0.2605 0.07515 0.07205 -0.0420 0.9838 0.0186
-5.750 -0.2491 0.06969 0.06660 -0.0432 0.9796 0.0171
-5.500 -0.2254 0.06499 0.06186 -0.0481 0.9734 0.0142
-5.250 -0.1917 0.06007 0.05684 -0.0554 0.9678 0.0149
-5.000 -0.1606 0.05549 0.05217 -0.0612 0.9599 0.0157
-4.750 -0.1234 0.05024 0.04678 -0.0679 0.9547 0.0146
-4.500 -0.0926 0.04541 0.04180 -0.0722 0.9458 0.0137
-4.250 -0.0557 0.04006 0.03621 -0.0772 0.9404 0.0132
-4.000 -0.0285 0.03539 0.03128 -0.0789 0.9307 0.0128
-3.750 0.0005 0.03019 0.02573 -0.0803 0.9226 0.0126
-3.500 0.0260 0.02339 0.01829 -0.0801 0.9144 0.0125
-3.250 0.0484 0.01885 0.01297 -0.0787 0.9045 0.0131
-3.000 0.0775 0.01691 0.01054 -0.0785 0.8955 0.0155
-2.750 0.1071 0.01507 0.00821 -0.0783 0.8856 0.0172
-2.500 0.1350 0.01393 0.00675 -0.0778 0.8742 0.0188
-2.250 0.1632 0.01305 0.00572 -0.0776 0.8631 0.0258
-2.000 0.1918 0.01232 0.00486 -0.0774 0.8516 0.0355
-1.750 0.2205 0.01190 0.00434 -0.0773 0.8391 0.0511
-1.500 0.2494 0.01180 0.00423 -0.0773 0.8255 0.0692
-1.250 0.2779 0.01170 0.00401 -0.0771 0.8103 0.0832
-1.000 0.3057 0.01155 0.00370 -0.0768 0.7935 0.0938
-0.750 0.3330 0.01142 0.00342 -0.0765 0.7740 0.1026
-0.500 0.3591 0.01130 0.00316 -0.0759 0.7509 0.1104
-0.250 0.3851 0.01128 0.00299 -0.0752 0.7257 0.1209
0.000 0.4099 0.01121 0.00280 -0.0745 0.7008 0.1301
0.250 0.4344 0.01118 0.00264 -0.0736 0.6758 0.1361
0.500 0.4585 0.01117 0.00252 -0.0727 0.6512 0.1416
0.750 0.4823 0.01123 0.00242 -0.0717 0.6258 0.1480
1.000 0.5057 0.01129 0.00238 -0.0707 0.5987 0.1575
1.250 0.5289 0.01136 0.00236 -0.0697 0.5697 0.1722
1.500 0.5517 0.01142 0.00236 -0.0686 0.5398 0.2010
2.000 0.6749 0.01046 0.00259 -0.0846 0.4634 1.0000
2.500 0.7217 0.01097 0.00284 -0.0829 0.4307 1.0000
2.750 0.7454 0.01121 0.00300 -0.0821 0.4179 1.0000
3.000 0.7693 0.01143 0.00319 -0.0813 0.4076 1.0000
3.250 0.7931 0.01168 0.00338 -0.0805 0.3984 1.0000
3.500 0.8173 0.01190 0.00358 -0.0798 0.3896 1.0000
3.750 0.8413 0.01213 0.00381 -0.0791 0.3824 1.0000
4.000 0.8655 0.01235 0.00407 -0.0784 0.3749 1.0000
4.250 0.8894 0.01259 0.00433 -0.0777 0.3676 1.0000
4.500 0.9125 0.01287 0.00459 -0.0768 0.3551 1.0000
4.750 0.9345 0.01317 0.00483 -0.0758 0.3299 1.0000
5.000 0.9564 0.01349 0.00512 -0.0748 0.3023 1.0000
5.250 0.9772 0.01392 0.00540 -0.0736 0.2569 1.0000
5.500 0.9883 0.01536 0.00611 -0.0712 0.1507 1.0000
5.750 0.9957 0.01733 0.00731 -0.0681 0.0215 1.0000
6.000 1.0161 0.01791 0.00803 -0.0667 0.0162 1.0000
6.250 1.0360 0.01855 0.00887 -0.0651 0.0144 1.0000
6.500 1.0531 0.01947 0.01001 -0.0632 0.0118 1.0000
6.750 1.0686 0.02053 0.01130 -0.0610 0.0106 1.0000
7.000 1.0847 0.02145 0.01240 -0.0590 0.0102 1.0000
7.250 1.0982 0.02250 0.01362 -0.0566 0.0098 1.0000
7.500 1.1095 0.02366 0.01490 -0.0539 0.0092 1.0000
7.750 1.1198 0.02484 0.01618 -0.0512 0.0087 1.0000
8.000 1.1291 0.02606 0.01745 -0.0485 0.0077 1.0000
8.250 1.1351 0.02755 0.01905 -0.0453 0.0072 1.0000
8.500 1.1397 0.02914 0.02068 -0.0419 0.0070 1.0000
8.750 1.1466 0.03132 0.02284 -0.0390 0.0068 1.0000
9.000 1.1626 0.03336 0.02491 -0.0375 0.0067 1.0000
9.250 1.1890 0.03635 0.02795 -0.0378 0.0066 1.0000
9.500 1.2079 0.03865 0.03042 -0.0367 0.0066 1.0000
9.750 1.2260 0.04133 0.03328 -0.0356 0.0066 1.0000
10.250 1.2508 0.04605 0.03843 -0.0320 0.0067 1.0000
10.500 1.2568 0.04890 0.04155 -0.0295 0.0067 1.0000
10.750 1.2335 0.04919 0.04219 -0.0229 0.0062 1.0000
11.000 1.2274 0.05096 0.04424 -0.0196 0.0059 1.0000
11.250 1.2230 0.05378 0.04729 -0.0172 0.0060 1.0000
11.500 1.2144 0.05624 0.05002 -0.0150 0.0057 1.0000
11.750 1.2056 0.05936 0.05338 -0.0134 0.0056 1.0000
12.000 1.1962 0.06292 0.05715 -0.0124 0.0057 1.0000
12.250 1.1833 0.06671 0.06118 -0.0121 0.0056 1.0000
12.500 1.1694 0.07095 0.06565 -0.0125 0.0054 1.0000
12.750 1.1561 0.07561 0.07050 -0.0135 0.0055 1.0000
13.000 1.1416 0.08076 0.07584 -0.0151 0.0056 1.0000
13.250 1.1253 0.08647 0.08175 -0.0176 0.0054 1.0000
13.500 1.1116 0.09218 0.08761 -0.0202 0.0056 1.0000
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Polar data table (+)
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