USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 50,000 Max Cl/Cd: 36.13 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa50-il-50000-n5.txt Download as CSV file: xf-usa50-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 50 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4931 0.10418 0.09738 -0.0195 1.0000 0.0870
-8.250 -0.5040 0.10172 0.09506 -0.0226 1.0000 0.0877
-8.000 -0.5091 0.09842 0.09185 -0.0255 1.0000 0.0879
-7.750 -0.4873 0.09130 0.08466 -0.0226 1.0000 0.0584
-7.500 -0.4871 0.08628 0.07969 -0.0256 1.0000 0.0459
-7.250 -0.4851 0.08242 0.07587 -0.0265 1.0000 0.0447
-6.750 -0.4814 0.07223 0.06553 -0.0338 1.0000 0.0373
-6.500 -0.4759 0.06822 0.06148 -0.0340 1.0000 0.0371
-6.250 -0.4693 0.06422 0.05740 -0.0343 1.0000 0.0369
-6.000 -0.4614 0.06021 0.05325 -0.0345 1.0000 0.0367
-5.750 -0.4519 0.05624 0.04911 -0.0345 1.0000 0.0365
-5.500 -0.4409 0.05233 0.04498 -0.0343 1.0000 0.0361
-5.250 -0.4281 0.04857 0.04097 -0.0338 1.0000 0.0356
-5.000 -0.4137 0.04486 0.03693 -0.0331 1.0000 0.0351
-4.750 -0.3974 0.04134 0.03300 -0.0323 1.0000 0.0348
-4.500 -0.3794 0.03789 0.02907 -0.0312 1.0000 0.0348
-4.250 -0.3597 0.03480 0.02547 -0.0299 1.0000 0.0352
-4.000 -0.3378 0.03202 0.02199 -0.0285 1.0000 0.0370
-3.750 -0.3173 0.02987 0.01963 -0.0275 1.0000 0.0410
-3.500 -0.2937 0.02770 0.01703 -0.0262 1.0000 0.0438
-3.250 -0.2686 0.02564 0.01450 -0.0248 1.0000 0.0470
-3.000 -0.2454 0.02416 0.01289 -0.0237 1.0000 0.0565
-2.750 -0.2197 0.02255 0.01105 -0.0227 1.0000 0.0685
-2.500 -0.1947 0.02114 0.00944 -0.0219 1.0000 0.1014
-2.250 -0.1710 0.01971 0.00821 -0.0214 1.0000 0.1610
-2.000 -0.1480 0.01798 0.00734 -0.0210 1.0000 0.3055
-1.750 -0.0813 0.01580 0.00672 -0.0269 1.0000 1.0000
-1.500 -0.0613 0.01582 0.00629 -0.0255 1.0000 1.0000
-1.250 -0.0414 0.01587 0.00599 -0.0243 1.0000 1.0000
-1.000 -0.0214 0.01594 0.00577 -0.0230 1.0000 1.0000
-0.750 -0.0014 0.01605 0.00558 -0.0218 1.0000 1.0000
-0.500 0.0183 0.01618 0.00550 -0.0207 1.0000 1.0000
-0.250 0.0379 0.01634 0.00549 -0.0195 1.0000 1.0000
0.000 0.0572 0.01654 0.00553 -0.0184 1.0000 1.0000
0.250 0.0760 0.01678 0.00565 -0.0173 1.0000 1.0000
0.500 0.1006 0.01708 0.00583 -0.0176 0.9969 1.0000
0.750 0.1458 0.01744 0.00611 -0.0219 0.9825 1.0000
1.000 0.1901 0.01774 0.00637 -0.0259 0.9678 1.0000
1.250 0.2301 0.01793 0.00656 -0.0290 0.9501 1.0000
1.500 0.2716 0.01810 0.00676 -0.0321 0.9328 1.0000
1.750 0.3150 0.01820 0.00696 -0.0354 0.9158 1.0000
2.000 0.3568 0.01819 0.00707 -0.0381 0.8949 1.0000
2.250 0.4029 0.01808 0.00712 -0.0414 0.8740 1.0000
2.500 0.4423 0.01803 0.00727 -0.0433 0.8490 1.0000
2.750 0.4798 0.01802 0.00744 -0.0448 0.8241 1.0000
3.000 0.5169 0.01803 0.00764 -0.0461 0.7989 1.0000
3.250 0.5533 0.01808 0.00793 -0.0471 0.7733 1.0000
3.500 0.5877 0.01818 0.00824 -0.0477 0.7468 1.0000
3.750 0.6184 0.01839 0.00866 -0.0476 0.7197 1.0000
4.000 0.6470 0.01867 0.00924 -0.0472 0.6922 1.0000
4.250 0.6741 0.01902 0.00986 -0.0464 0.6645 1.0000
4.500 0.6947 0.01923 0.01005 -0.0433 0.6043 1.0000
4.750 0.7067 0.01967 0.01014 -0.0387 0.5096 1.0000
5.000 0.7101 0.02075 0.01025 -0.0334 0.3189 1.0000
5.250 0.7152 0.02432 0.01191 -0.0310 0.0685 1.0000
5.500 0.7336 0.02610 0.01383 -0.0293 0.0489 1.0000
5.750 0.7510 0.02768 0.01556 -0.0275 0.0399 1.0000
6.000 0.7671 0.02930 0.01731 -0.0256 0.0350 1.0000
6.250 0.7839 0.03103 0.01930 -0.0234 0.0330 1.0000
6.500 0.8032 0.03293 0.02148 -0.0215 0.0313 1.0000
6.750 0.8244 0.03500 0.02378 -0.0200 0.0288 1.0000
7.000 0.8452 0.03761 0.02645 -0.0190 0.0266 1.0000
7.250 0.8685 0.04041 0.02973 -0.0176 0.0257 1.0000
7.500 0.8880 0.04358 0.03346 -0.0161 0.0255 1.0000
7.750 0.9031 0.04696 0.03732 -0.0143 0.0255 1.0000
8.000 0.9140 0.05049 0.04130 -0.0123 0.0256 1.0000
8.250 0.9209 0.05414 0.04538 -0.0101 0.0257 1.0000
8.500 0.9242 0.05784 0.04946 -0.0080 0.0259 1.0000
8.750 0.9239 0.06165 0.05361 -0.0060 0.0260 1.0000
9.000 0.9207 0.06542 0.05765 -0.0041 0.0262 1.0000
9.250 0.9144 0.06919 0.06166 -0.0023 0.0264 1.0000
9.500 0.9038 0.07292 0.06558 -0.0006 0.0265 1.0000
9.750 0.8908 0.07651 0.06931 0.0008 0.0266 1.0000
10.000 0.8763 0.08048 0.07346 0.0011 0.0267 1.0000
10.250 0.8508 0.08538 0.07860 -0.0015 0.0274 1.0000
10.500 0.8306 0.09191 0.08523 -0.0062 0.0277 1.0000
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