USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 49 AIRFOIL (usa49-il) Reynolds number: 100,000 Max Cl/Cd: 42.13 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa49-il-100000-n5.txt Download as CSV file: xf-usa49-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 49 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4358 0.08726 0.08276 -0.0107 1.0000 0.0495
-8.250 -0.4380 0.08294 0.07848 -0.0121 1.0000 0.0492
-8.000 -0.4413 0.07844 0.07402 -0.0137 1.0000 0.0482
-7.500 -0.4573 0.06613 0.06178 -0.0193 1.0000 0.0264
-7.000 -0.5287 0.06644 0.06162 -0.0261 1.0000 0.0217
-6.750 -0.5209 0.06179 0.05688 -0.0280 1.0000 0.0212
-6.500 -0.5108 0.05682 0.05176 -0.0299 1.0000 0.0209
-6.250 -0.4988 0.05182 0.04654 -0.0314 1.0000 0.0209
-6.000 -0.4850 0.04698 0.04141 -0.0322 1.0000 0.0209
-5.750 -0.4696 0.04249 0.03656 -0.0325 1.0000 0.0211
-5.500 -0.4530 0.03838 0.03204 -0.0322 1.0000 0.0211
-5.250 -0.4350 0.03462 0.02783 -0.0316 1.0000 0.0208
-5.000 -0.4155 0.03123 0.02395 -0.0307 1.0000 0.0206
-4.750 -0.3944 0.02826 0.02043 -0.0296 1.0000 0.0206
-4.500 -0.3721 0.02572 0.01742 -0.0285 1.0000 0.0207
-4.250 -0.3489 0.02359 0.01486 -0.0274 1.0000 0.0210
-4.000 -0.3250 0.02197 0.01290 -0.0263 1.0000 0.0220
-3.750 -0.3019 0.02015 0.01089 -0.0253 1.0000 0.0239
-3.500 -0.2788 0.01877 0.00941 -0.0244 1.0000 0.0256
-3.250 -0.2555 0.01766 0.00817 -0.0234 1.0000 0.0274
-3.000 -0.2320 0.01677 0.00714 -0.0224 1.0000 0.0302
-2.750 -0.2084 0.01595 0.00624 -0.0217 1.0000 0.0363
-2.500 -0.1842 0.01522 0.00545 -0.0209 1.0000 0.0452
-2.250 -0.1601 0.01426 0.00496 -0.0204 1.0000 0.1380
-2.000 -0.1363 0.01320 0.00465 -0.0203 1.0000 0.2993
-1.750 -0.1208 0.01135 0.00474 -0.0170 1.0000 0.7646
-1.500 -0.0737 0.01115 0.00459 -0.0197 1.0000 0.9709
-1.250 -0.0447 0.01116 0.00435 -0.0203 1.0000 1.0000
-1.000 -0.0243 0.01123 0.00424 -0.0192 1.0000 1.0000
-0.750 0.0166 0.01130 0.00410 -0.0223 0.9911 1.0000
-0.500 0.0581 0.01136 0.00398 -0.0254 0.9816 1.0000
-0.250 0.0999 0.01140 0.00390 -0.0285 0.9713 1.0000
0.000 0.1413 0.01143 0.00382 -0.0314 0.9592 1.0000
0.250 0.1810 0.01142 0.00375 -0.0339 0.9440 1.0000
0.500 0.2186 0.01140 0.00368 -0.0357 0.9267 1.0000
0.750 0.2506 0.01139 0.00363 -0.0364 0.9057 1.0000
1.000 0.2823 0.01137 0.00358 -0.0369 0.8847 1.0000
1.250 0.3110 0.01137 0.00356 -0.0367 0.8602 1.0000
1.500 0.3389 0.01138 0.00355 -0.0363 0.8339 1.0000
1.750 0.3660 0.01139 0.00351 -0.0356 0.8026 1.0000
2.000 0.3916 0.01144 0.00343 -0.0344 0.7600 1.0000
2.250 0.4161 0.01158 0.00340 -0.0330 0.7102 1.0000
2.500 0.4405 0.01181 0.00343 -0.0319 0.6620 1.0000
2.750 0.4646 0.01209 0.00352 -0.0307 0.6139 1.0000
3.000 0.4882 0.01245 0.00367 -0.0297 0.5606 1.0000
3.250 0.5128 0.01279 0.00387 -0.0289 0.5195 1.0000
3.500 0.5378 0.01311 0.00412 -0.0283 0.4844 1.0000
3.750 0.5626 0.01346 0.00439 -0.0277 0.4420 1.0000
4.000 0.5865 0.01392 0.00469 -0.0270 0.3809 1.0000
4.250 0.6098 0.01453 0.00502 -0.0263 0.3150 1.0000
4.500 0.6329 0.01527 0.00546 -0.0257 0.2398 1.0000
4.750 0.6548 0.01632 0.00604 -0.0252 0.1424 1.0000
5.000 0.6775 0.01730 0.00677 -0.0246 0.0966 1.0000
5.250 0.7015 0.01803 0.00752 -0.0241 0.0879 1.0000
5.500 0.7256 0.01872 0.00834 -0.0235 0.0832 1.0000
5.750 0.7490 0.01950 0.00927 -0.0229 0.0797 1.0000
6.000 0.7711 0.02049 0.01042 -0.0220 0.0770 1.0000
6.250 0.7940 0.02135 0.01148 -0.0212 0.0740 1.0000
6.500 0.8155 0.02244 0.01268 -0.0205 0.0666 1.0000
6.750 0.8367 0.02361 0.01400 -0.0198 0.0556 1.0000
7.000 0.8573 0.02499 0.01555 -0.0189 0.0455 1.0000
7.250 0.8782 0.02639 0.01712 -0.0180 0.0357 1.0000
7.500 0.8994 0.02762 0.01848 -0.0172 0.0291 1.0000
7.750 0.9208 0.02938 0.02054 -0.0161 0.0240 1.0000
8.000 0.9405 0.03093 0.02232 -0.0152 0.0216 1.0000
8.250 0.9591 0.03361 0.02532 -0.0140 0.0197 1.0000
8.500 0.9760 0.03681 0.02903 -0.0125 0.0177 1.0000
8.750 0.9884 0.04062 0.03336 -0.0108 0.0169 1.0000
9.000 0.9959 0.04476 0.03802 -0.0090 0.0163 1.0000
9.250 0.9979 0.04915 0.04292 -0.0072 0.0160 1.0000
9.500 0.9945 0.05365 0.04787 -0.0056 0.0159 1.0000
9.750 0.9855 0.05819 0.05278 -0.0041 0.0158 1.0000
10.000 0.9706 0.06249 0.05736 -0.0028 0.0159 1.0000
10.250 0.9518 0.06705 0.06214 -0.0029 0.0159 1.0000
10.500 0.9318 0.07263 0.06790 -0.0055 0.0161 1.0000
10.750 0.9112 0.07989 0.07531 -0.0111 0.0163 1.0000
11.000 0.8908 0.08919 0.08471 -0.0187 0.0167 1.0000
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