USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 50 AIRFOIL (usa50-il) Reynolds number: 200,000 Max Cl/Cd: 60.51 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa50-il-200000-n5.txt Download as CSV file: xf-usa50-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4768 0.09164 0.08820 -0.0225 1.0000 0.0188
-8.000 -0.4793 0.08813 0.08474 -0.0238 1.0000 0.0188
-7.750 -0.4834 0.08474 0.08140 -0.0250 1.0000 0.0189
-7.500 -0.4831 0.08056 0.07725 -0.0273 1.0000 0.0189
-7.250 -0.4130 0.06156 0.05850 -0.0253 1.0000 0.0161
-7.000 -0.4238 0.05797 0.05495 -0.0259 1.0000 0.0154
-6.500 -0.4837 0.06404 0.06072 -0.0294 1.0000 0.0157
-6.250 -0.4790 0.05919 0.05580 -0.0307 1.0000 0.0139
-6.000 -0.4680 0.05106 0.04742 -0.0325 1.0000 0.0104
-5.750 -0.4552 0.04845 0.04468 -0.0317 1.0000 0.0096
-5.500 -0.4479 0.04379 0.03981 -0.0311 1.0000 0.0092
-5.250 -0.4370 0.03953 0.03530 -0.0301 1.0000 0.0089
-5.000 -0.4244 0.03544 0.03090 -0.0288 1.0000 0.0086
-4.750 -0.4098 0.03146 0.02656 -0.0271 1.0000 0.0084
-4.500 -0.3930 0.02762 0.02228 -0.0252 1.0000 0.0087
-4.250 -0.3699 0.02499 0.01921 -0.0242 0.9992 0.0095
-4.000 -0.3381 0.02197 0.01558 -0.0250 0.9962 0.0101
-3.750 -0.3073 0.01870 0.01167 -0.0253 0.9933 0.0105
-3.500 -0.2755 0.01652 0.00910 -0.0258 0.9901 0.0110
-3.250 -0.2428 0.01505 0.00740 -0.0267 0.9870 0.0122
-3.000 -0.2126 0.01415 0.00638 -0.0272 0.9826 0.0144
-2.750 -0.1801 0.01335 0.00540 -0.0281 0.9789 0.0177
-2.500 -0.1508 0.01262 0.00456 -0.0283 0.9737 0.0245
-2.250 -0.1181 0.01213 0.00417 -0.0294 0.9691 0.0620
-2.000 -0.0883 0.01180 0.00387 -0.0299 0.9625 0.0898
-1.750 -0.0558 0.01135 0.00352 -0.0311 0.9569 0.1275
-1.500 -0.0298 0.01054 0.00324 -0.0312 0.9490 0.2796
-1.250 -0.0028 0.00958 0.00309 -0.0314 0.9428 0.5100
-1.000 0.0177 0.00871 0.00302 -0.0294 0.9339 0.7293
-0.750 0.1161 0.00851 0.00315 -0.0445 0.9451 1.0000
-0.500 0.1486 0.00843 0.00297 -0.0455 0.9345 1.0000
-0.250 0.1849 0.00835 0.00278 -0.0473 0.9249 1.0000
0.000 0.2267 0.00825 0.00260 -0.0502 0.9133 1.0000
0.250 0.2736 0.00811 0.00239 -0.0542 0.8940 1.0000
0.500 0.3258 0.00796 0.00214 -0.0594 0.8630 1.0000
0.750 0.3707 0.00791 0.00197 -0.0629 0.8212 1.0000
1.000 0.4055 0.00798 0.00188 -0.0642 0.7760 1.0000
1.250 0.4329 0.00815 0.00186 -0.0639 0.7346 1.0000
1.500 0.4570 0.00835 0.00191 -0.0630 0.6996 1.0000
1.750 0.4801 0.00856 0.00202 -0.0619 0.6683 1.0000
2.000 0.5031 0.00876 0.00214 -0.0608 0.6412 1.0000
2.250 0.5257 0.00897 0.00228 -0.0596 0.6154 1.0000
2.500 0.5478 0.00921 0.00243 -0.0584 0.5880 1.0000
2.750 0.5697 0.00948 0.00264 -0.0571 0.5617 1.0000
3.000 0.5912 0.00977 0.00285 -0.0557 0.5319 1.0000
3.250 0.6095 0.01019 0.00305 -0.0537 0.4697 1.0000
3.500 0.6273 0.01066 0.00322 -0.0516 0.3788 1.0000
3.750 0.6393 0.01181 0.00363 -0.0489 0.2341 1.0000
4.000 0.6476 0.01373 0.00443 -0.0458 0.0257 1.0000
4.250 0.6686 0.01436 0.00509 -0.0443 0.0175 1.0000
4.500 0.6900 0.01493 0.00584 -0.0429 0.0140 1.0000
4.750 0.7102 0.01565 0.00675 -0.0412 0.0123 1.0000
5.000 0.7285 0.01659 0.00786 -0.0393 0.0112 1.0000
5.250 0.7418 0.01818 0.00963 -0.0365 0.0100 1.0000
5.500 0.7615 0.01905 0.01060 -0.0349 0.0089 1.0000
5.750 0.7803 0.02025 0.01190 -0.0331 0.0084 1.0000
6.000 0.7998 0.02172 0.01349 -0.0313 0.0082 1.0000
6.250 0.8204 0.02339 0.01533 -0.0298 0.0080 1.0000
6.500 0.8414 0.02533 0.01750 -0.0282 0.0079 1.0000
6.750 0.8614 0.02754 0.02001 -0.0265 0.0079 1.0000
7.000 0.8794 0.02997 0.02280 -0.0246 0.0080 1.0000
7.250 0.8943 0.03283 0.02613 -0.0222 0.0081 1.0000
7.500 0.9060 0.03590 0.02960 -0.0196 0.0084 1.0000
7.750 0.9183 0.03809 0.03214 -0.0172 0.0075 1.0000
8.000 0.9279 0.04027 0.03458 -0.0150 0.0068 1.0000
8.250 0.9330 0.04323 0.03787 -0.0123 0.0066 1.0000
8.500 0.9344 0.04653 0.04150 -0.0094 0.0066 1.0000
8.750 0.9341 0.04927 0.04444 -0.0069 0.0063 1.0000
9.000 0.9220 0.05432 0.04995 -0.0030 0.0070 1.0000
9.250 0.9075 0.05803 0.05389 0.0005 0.0074 1.0000
9.500 0.8903 0.06150 0.05753 0.0032 0.0073 1.0000
9.750 0.8727 0.06538 0.06157 0.0040 0.0076 1.0000
10.000 0.8544 0.07007 0.06641 0.0027 0.0076 1.0000
10.250 0.8381 0.07559 0.07204 -0.0008 0.0076 1.0000
10.500 0.8221 0.08280 0.07934 -0.0063 0.0076 1.0000
10.750 0.8091 0.09123 0.08783 -0.0126 0.0079 1.0000
11.000 0.8001 0.10051 0.09711 -0.0182 0.0079 1.0000
11.250 0.7920 0.10797 0.10453 -0.0220 0.0084 1.0000
11.500 0.6730 0.10296 0.09979 -0.0157 0.0102 1.0000
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