USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 500,000 Max Cl/Cd: 80.42 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa50-il-500000.txt Download as CSV file: xf-usa50-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.5022 0.08315 0.08097 -0.0178 1.0000 0.0135
-7.750 -0.5098 0.07994 0.07781 -0.0185 1.0000 0.0136
-7.500 -0.5157 0.07621 0.07410 -0.0202 1.0000 0.0138
-7.250 -0.5163 0.07195 0.06985 -0.0229 1.0000 0.0138
-7.000 -0.5151 0.06746 0.06534 -0.0253 1.0000 0.0140
-6.750 -0.5117 0.06286 0.06069 -0.0274 1.0000 0.0146
-6.500 -0.5065 0.05838 0.05613 -0.0286 1.0000 0.0148
-6.250 -0.4983 0.05383 0.05146 -0.0294 1.0000 0.0157
-6.000 -0.4797 0.05079 0.04826 -0.0291 1.0000 0.0170
-5.750 -0.4692 0.04713 0.04443 -0.0282 1.0000 0.0172
-5.500 -0.4594 0.04345 0.04056 -0.0268 1.0000 0.0172
-5.250 -0.4490 0.03987 0.03676 -0.0252 1.0000 0.0173
-4.250 -0.3787 0.02156 0.01684 -0.0217 0.9961 0.0177
-4.000 -0.3474 0.01878 0.01365 -0.0225 0.9942 0.0178
-3.750 -0.3200 0.01546 0.00992 -0.0220 0.9922 0.0160
-3.500 -0.2900 0.01350 0.00767 -0.0221 0.9900 0.0157
-3.250 -0.2583 0.01226 0.00624 -0.0227 0.9876 0.0162
-3.000 -0.2242 0.01145 0.00531 -0.0239 0.9852 0.0175
-2.750 -0.1901 0.01039 0.00412 -0.0252 0.9831 0.0214
-2.500 -0.1599 0.00994 0.00360 -0.0256 0.9785 0.0261
-2.250 -0.1265 0.00941 0.00318 -0.0266 0.9747 0.0599
-2.000 -0.0896 0.00924 0.00302 -0.0287 0.9719 0.0789
-1.750 -0.0557 0.00893 0.00276 -0.0301 0.9682 0.0955
-1.500 -0.0255 0.00843 0.00249 -0.0307 0.9623 0.1554
-1.250 0.0023 0.00702 0.00235 -0.0315 0.9583 0.5199
-1.000 0.0228 0.00616 0.00224 -0.0297 0.9502 0.7304
-0.750 0.1256 0.00587 0.00254 -0.0459 0.9614 0.9919
-0.500 0.1764 0.00578 0.00238 -0.0510 0.9574 1.0000
-0.250 0.2118 0.00559 0.00213 -0.0524 0.9441 1.0000
0.000 0.2576 0.00543 0.00194 -0.0563 0.9307 1.0000
0.250 0.3148 0.00530 0.00174 -0.0628 0.9131 1.0000
0.500 0.3712 0.00523 0.00156 -0.0690 0.8815 1.0000
0.750 0.4058 0.00531 0.00146 -0.0703 0.8320 1.0000
1.000 0.4286 0.00550 0.00142 -0.0690 0.7722 1.0000
1.250 0.4477 0.00576 0.00141 -0.0670 0.7177 1.0000
1.500 0.4670 0.00602 0.00144 -0.0651 0.6697 1.0000
1.750 0.4872 0.00624 0.00149 -0.0634 0.6276 1.0000
2.000 0.5081 0.00645 0.00155 -0.0619 0.5932 1.0000
2.250 0.5286 0.00671 0.00166 -0.0603 0.5561 1.0000
2.500 0.5504 0.00691 0.00176 -0.0590 0.5295 1.0000
2.750 0.5726 0.00712 0.00187 -0.0578 0.5015 1.0000
3.000 0.5935 0.00738 0.00196 -0.0564 0.4501 1.0000
3.250 0.6125 0.00782 0.00207 -0.0547 0.3634 1.0000
3.500 0.6286 0.00864 0.00237 -0.0525 0.2521 1.0000
3.750 0.6402 0.01003 0.00293 -0.0498 0.0741 1.0000
4.000 0.6590 0.01078 0.00341 -0.0479 0.0207 1.0000
4.250 0.6815 0.01114 0.00388 -0.0467 0.0185 1.0000
4.500 0.7031 0.01162 0.00447 -0.0452 0.0170 1.0000
4.750 0.7237 0.01220 0.00517 -0.0436 0.0161 1.0000
5.000 0.7430 0.01291 0.00598 -0.0417 0.0152 1.0000
5.250 0.7620 0.01364 0.00679 -0.0399 0.0135 1.0000
5.500 0.7792 0.01462 0.00787 -0.0377 0.0128 1.0000
5.750 0.7942 0.01614 0.00949 -0.0351 0.0119 1.0000
6.000 0.8123 0.01772 0.01118 -0.0330 0.0117 1.0000
6.250 0.8335 0.01895 0.01252 -0.0314 0.0121 1.0000
6.500 0.8572 0.02270 0.01672 -0.0290 0.0177 1.0000
6.750 0.8491 0.01371 0.00842 -0.0224 0.0249 1.0000
7.000 0.8654 0.01518 0.01005 -0.0207 0.0226 1.0000
7.250 0.8803 0.01721 0.01218 -0.0191 0.0211 1.0000
19.000 0.7050 0.22203 0.22006 -0.0746 0.0072 1.0000
19.250 0.7096 0.22545 0.22349 -0.0763 0.0071 1.0000
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