USA 25 AIRFOIL (usa25-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 25 AIRFOIL (usa25-il) Reynolds number: 100,000 Max Cl/Cd: 52.57 at α=9.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa25-il-100000.txt Download as CSV file: xf-usa25-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: USA 25 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3412 0.09638 0.09193 -0.0191 1.0000 0.0845
-7.000 -0.3532 0.09496 0.09060 -0.0175 1.0000 0.0859
-6.750 -0.3636 0.09379 0.08949 -0.0178 1.0000 0.0878
-6.500 -0.3748 0.09407 0.08978 -0.0205 1.0000 0.0890
-6.250 -0.3774 0.09421 0.08982 -0.0236 1.0000 0.0895
-6.000 -0.3756 0.08750 0.08327 -0.0194 1.0000 0.0903
-5.750 -0.3729 0.08363 0.07945 -0.0161 1.0000 0.0915
-5.500 -0.3706 0.08079 0.07665 -0.0140 1.0000 0.0930
-5.250 -0.3677 0.07833 0.07420 -0.0129 1.0000 0.0949
-5.000 -0.3633 0.07599 0.07187 -0.0126 1.0000 0.0979
-4.500 -0.3417 0.07179 0.06744 -0.0173 1.0000 0.1048
-4.250 -0.3248 0.06759 0.06330 -0.0174 0.9965 0.1069
-4.000 -0.2858 0.06379 0.05940 -0.0228 0.9879 0.1118
-3.750 -0.2309 0.06076 0.05599 -0.0327 0.9774 0.1195
-3.500 -0.1975 0.05631 0.05157 -0.0359 0.9699 0.1231
-3.250 -0.1472 0.05404 0.04894 -0.0427 0.9582 0.1345
-3.000 -0.1169 0.05000 0.04493 -0.0448 0.9474 0.1381
-2.750 -0.0666 0.04805 0.04259 -0.0504 0.9352 0.1495
-2.500 -0.0277 0.04391 0.03848 -0.0537 0.9265 0.1541
-2.250 0.0176 0.04150 0.03577 -0.0576 0.9140 0.1656
-2.000 0.0585 0.03924 0.03331 -0.0603 0.9002 0.1798
-1.500 0.1502 0.02967 0.02261 -0.0647 0.8740 0.1067
-1.250 0.1957 0.02614 0.01858 -0.0667 0.8598 0.1000
-1.000 0.2458 0.02309 0.01504 -0.0695 0.8432 0.1001
-0.750 0.2961 0.02103 0.01267 -0.0726 0.8183 0.1078
-0.500 0.3486 0.01924 0.01046 -0.0761 0.7859 0.1194
-0.250 0.3947 0.01826 0.00924 -0.0787 0.7485 0.1466
0.000 0.4277 0.01805 0.00879 -0.0792 0.7128 0.1797
0.250 0.4582 0.01812 0.00867 -0.0796 0.6869 0.2053
0.500 0.4869 0.01806 0.00842 -0.0796 0.6662 0.2205
0.750 0.5150 0.01806 0.00828 -0.0796 0.6502 0.2392
1.000 0.5437 0.01798 0.00811 -0.0797 0.6369 0.2556
1.250 0.5745 0.01792 0.00790 -0.0800 0.6255 0.2669
1.500 0.6047 0.01781 0.00778 -0.0804 0.6150 0.2803
1.750 0.6322 0.01774 0.00771 -0.0802 0.6063 0.2977
2.000 0.6569 0.01767 0.00762 -0.0796 0.5984 0.3247
2.250 0.6780 0.01743 0.00764 -0.0784 0.5910 0.3834
2.500 0.7975 0.01686 0.00797 -0.0969 0.5798 1.0000
2.750 0.8188 0.01723 0.00830 -0.0956 0.5727 1.0000
3.000 0.8432 0.01765 0.00857 -0.0950 0.5669 1.0000
3.250 0.8640 0.01807 0.00901 -0.0937 0.5607 1.0000
3.500 0.8862 0.01849 0.00937 -0.0926 0.5538 1.0000
3.750 0.9075 0.01892 0.00976 -0.0913 0.5462 1.0000
4.000 0.9289 0.01930 0.01008 -0.0900 0.5375 1.0000
4.250 0.9485 0.01972 0.01053 -0.0884 0.5289 1.0000
4.500 0.9714 0.02013 0.01087 -0.0874 0.5212 1.0000
4.750 0.9907 0.02063 0.01144 -0.0859 0.5145 1.0000
5.000 1.0123 0.02110 0.01194 -0.0847 0.5077 1.0000
5.250 1.0338 0.02163 0.01249 -0.0835 0.5013 1.0000
5.500 1.0530 0.02213 0.01309 -0.0820 0.4939 1.0000
5.750 1.0770 0.02268 0.01360 -0.0812 0.4876 1.0000
6.000 1.0934 0.02325 0.01437 -0.0792 0.4800 1.0000
6.250 1.1194 0.02377 0.01483 -0.0788 0.4732 1.0000
6.500 1.1332 0.02436 0.01564 -0.0763 0.4644 1.0000
6.750 1.1580 0.02487 0.01611 -0.0757 0.4568 1.0000
7.000 1.1730 0.02544 0.01687 -0.0734 0.4472 1.0000
7.250 1.1927 0.02594 0.01746 -0.0718 0.4376 1.0000
7.500 1.2152 0.02603 0.01750 -0.0705 0.4250 1.0000
7.750 1.2350 0.02600 0.01745 -0.0687 0.4114 1.0000
8.000 1.2490 0.02616 0.01776 -0.0660 0.3984 1.0000
8.250 1.2643 0.02611 0.01780 -0.0635 0.3844 1.0000
8.500 1.2786 0.02583 0.01758 -0.0607 0.3696 1.0000
8.750 1.2920 0.02565 0.01749 -0.0578 0.3556 1.0000
9.000 1.3053 0.02558 0.01753 -0.0550 0.3425 1.0000
9.250 1.3176 0.02553 0.01761 -0.0520 0.3292 1.0000
9.500 1.3260 0.02538 0.01760 -0.0484 0.3123 1.0000
9.750 1.3342 0.02538 0.01768 -0.0448 0.2946 1.0000
10.000 1.3417 0.02559 0.01791 -0.0412 0.2773 1.0000
10.250 1.3466 0.02603 0.01841 -0.0374 0.2578 1.0000
10.500 1.3460 0.02667 0.01896 -0.0328 0.2358 1.0000
10.750 1.3451 0.02763 0.01988 -0.0286 0.2128 1.0000
11.000 1.3409 0.02894 0.02111 -0.0243 0.1893 1.0000
11.250 1.3321 0.03071 0.02278 -0.0199 0.1605 1.0000
11.500 1.3192 0.03301 0.02498 -0.0158 0.1191 1.0000
11.750 1.2970 0.03624 0.02787 -0.0118 0.0859 1.0000
12.000 1.2801 0.03948 0.03105 -0.0091 0.0745 1.0000
12.250 1.2626 0.04320 0.03479 -0.0072 0.0662 1.0000
12.500 1.2444 0.04745 0.03907 -0.0062 0.0619 1.0000
12.750 1.2330 0.05139 0.04314 -0.0059 0.0579 1.0000
13.000 1.2205 0.05580 0.04766 -0.0063 0.0548 1.0000
13.250 1.2070 0.06055 0.05249 -0.0070 0.0527 1.0000
13.500 1.1942 0.06528 0.05728 -0.0078 0.0511 1.0000
13.750 1.1890 0.06904 0.06116 -0.0081 0.0494 1.0000
14.000 1.1852 0.07257 0.06481 -0.0083 0.0475 1.0000
14.250 1.1844 0.07562 0.06794 -0.0082 0.0462 1.0000
14.500 1.1849 0.07845 0.07082 -0.0079 0.0444 1.0000
14.750 1.1897 0.08054 0.07293 -0.0072 0.0432 1.0000
15.000 1.2084 0.08022 0.07244 -0.0038 0.0413 1.0000
15.250 1.2187 0.08204 0.07445 -0.0028 0.0404 1.0000
15.500 1.2250 0.08455 0.07713 -0.0023 0.0394 1.0000
15.750 1.2287 0.08750 0.08026 -0.0023 0.0384 1.0000
16.000 1.2319 0.09064 0.08357 -0.0024 0.0376 1.0000
16.250 1.2311 0.09441 0.08752 -0.0031 0.0369 1.0000
16.500 1.2302 0.09836 0.09168 -0.0037 0.0368 1.0000
16.750 1.2231 0.10332 0.09687 -0.0053 0.0368 1.0000
17.000 1.2115 0.10911 0.10290 -0.0077 0.0369 1.0000
17.250 1.1959 0.11573 0.10976 -0.0110 0.0371 1.0000
17.500 1.1760 0.12336 0.11764 -0.0151 0.0375 1.0000
17.750 1.1530 0.13195 0.12644 -0.0202 0.0379 1.0000
18.000 1.1269 0.14178 0.13648 -0.0264 0.0385 1.0000
18.250 1.0968 0.15335 0.14821 -0.0339 0.0393 1.0000
18.500 1.0676 0.16581 0.16077 -0.0416 0.0403 1.0000
18.750 1.0480 0.17658 0.17153 -0.0476 0.0411 1.0000
19.000 0.9806 0.21298 0.20770 -0.0664 0.0591 1.0000
19.250 0.9905 0.21640 0.21115 -0.0668 0.0610 1.0000
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