USA 25 AIRFOIL (usa25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 25 AIRFOIL (usa25-il) Reynolds number: 1,000,000 Max Cl/Cd: 106.57 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa25-il-1000000-n5.txt Download as CSV file: xf-usa25-il-1000000-n5.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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.250 -0.4116 0.15845 0.15671 -0.0157 1.0000 0.0045
-14.000 -0.4115 0.15389 0.15216 -0.0165 1.0000 0.0051
-13.750 -0.4124 0.14917 0.14745 -0.0174 1.0000 0.0053
-9.750 -0.2952 0.09099 0.08930 -0.0516 0.9866 0.0104
-8.750 -0.2702 0.07411 0.07242 -0.0667 0.9743 0.0131
-8.500 -0.2673 0.07115 0.06947 -0.0692 0.9669 0.0132
-8.000 -0.2337 0.06426 0.06253 -0.0784 0.9564 0.0140
-7.500 -0.2076 0.04785 0.04586 -0.0956 0.9402 0.0168
-7.250 -0.2533 0.02162 0.01852 -0.1006 0.9227 0.0197
-7.000 -0.2305 0.01859 0.01506 -0.1010 0.9069 0.0201
-6.750 -0.2024 0.01727 0.01340 -0.1016 0.8715 0.0204
-6.500 -0.1821 0.01663 0.01236 -0.1002 0.8206 0.0206
-6.250 -0.1627 0.01629 0.01175 -0.0985 0.7841 0.0208
-6.000 -0.1418 0.01600 0.01124 -0.0970 0.7537 0.0209
-5.750 -0.1235 0.01521 0.01015 -0.0952 0.7168 0.0210
-5.500 -0.1084 0.01415 0.00869 -0.0929 0.6627 0.0213
-5.250 -0.0913 0.01369 0.00786 -0.0908 0.5962 0.0215
-5.000 -0.0715 0.01333 0.00721 -0.0892 0.5404 0.0217
-4.750 -0.0498 0.01303 0.00666 -0.0880 0.4988 0.0218
-4.500 -0.0265 0.01270 0.00617 -0.0870 0.4751 0.0220
-4.000 0.0228 0.01209 0.00534 -0.0857 0.4509 0.0225
-3.750 0.0479 0.01185 0.00502 -0.0851 0.4428 0.0228
-3.500 0.0733 0.01155 0.00464 -0.0845 0.4366 0.0230
-3.250 0.0987 0.01129 0.00430 -0.0839 0.4309 0.0233
-3.000 0.1242 0.01104 0.00399 -0.0833 0.4264 0.0235
-2.750 0.1500 0.01082 0.00372 -0.0828 0.4223 0.0239
-2.500 0.1756 0.01062 0.00346 -0.0822 0.4179 0.0242
-2.250 0.2011 0.01046 0.00325 -0.0817 0.4138 0.0246
-2.000 0.2269 0.01030 0.00305 -0.0811 0.4107 0.0250
-1.750 0.2527 0.01015 0.00287 -0.0806 0.4081 0.0253
-1.500 0.2784 0.01001 0.00270 -0.0801 0.4054 0.0255
-1.250 0.3041 0.00990 0.00256 -0.0795 0.4027 0.0257
-1.000 0.3296 0.00980 0.00244 -0.0790 0.3999 0.0259
-0.500 0.3808 0.00963 0.00223 -0.0778 0.3949 0.0263
-0.250 0.4066 0.00955 0.00214 -0.0773 0.3929 0.0264
0.000 0.4322 0.00949 0.00207 -0.0768 0.3902 0.0266
0.250 0.4578 0.00944 0.00202 -0.0762 0.3875 0.0268
0.500 0.4829 0.00939 0.00193 -0.0756 0.3842 0.0274
0.750 0.5080 0.00936 0.00187 -0.0749 0.3812 0.0279
1.000 0.5337 0.00931 0.00182 -0.0744 0.3791 0.0288
1.250 0.5594 0.00928 0.00179 -0.0739 0.3765 0.0299
1.500 0.5849 0.00926 0.00177 -0.0733 0.3732 0.0313
1.750 0.6104 0.00926 0.00177 -0.0728 0.3704 0.0327
2.250 0.6604 0.00924 0.00182 -0.0715 0.3650 0.0641
2.500 0.6860 0.00923 0.00186 -0.0711 0.3630 0.0763
2.750 0.7116 0.00923 0.00191 -0.0706 0.3600 0.0862
3.000 0.7369 0.00925 0.00196 -0.0700 0.3564 0.0988
3.250 0.7616 0.00926 0.00204 -0.0694 0.3525 0.1257
3.500 0.7868 0.00929 0.00212 -0.0689 0.3488 0.1431
3.750 0.8123 0.00931 0.00219 -0.0684 0.3457 0.1572
4.000 0.8376 0.00935 0.00227 -0.0679 0.3420 0.1693
4.250 0.8625 0.00941 0.00236 -0.0673 0.3377 0.1857
4.500 0.8873 0.00946 0.00245 -0.0667 0.3289 0.2083
4.750 0.9113 0.00954 0.00255 -0.0660 0.3179 0.2380
5.000 0.9348 0.00963 0.00268 -0.0652 0.3038 0.2760
5.250 0.9443 0.00895 0.00283 -0.0618 0.2855 0.7482
5.500 0.9527 0.00894 0.00315 -0.0576 0.2452 0.9281
5.750 1.0003 0.00986 0.00378 -0.0628 0.1817 0.9687
6.000 1.0479 0.01050 0.00425 -0.0679 0.1514 0.9842
6.250 1.0844 0.01106 0.00465 -0.0703 0.1236 0.9933
6.750 1.1208 0.01312 0.00611 -0.0677 0.0155 1.0000
7.000 1.1391 0.01337 0.00638 -0.0659 0.0134 1.0000
7.250 1.1574 0.01365 0.00667 -0.0641 0.0122 1.0000
7.500 1.1754 0.01395 0.00699 -0.0622 0.0112 1.0000
7.750 1.1927 0.01429 0.00737 -0.0603 0.0099 1.0000
8.000 1.2109 0.01459 0.00769 -0.0585 0.0094 1.0000
8.250 1.2283 0.01491 0.00803 -0.0567 0.0089 1.0000
8.500 1.2437 0.01524 0.00839 -0.0544 0.0084 1.0000
8.750 1.2580 0.01560 0.00877 -0.0519 0.0079 1.0000
9.000 1.2715 0.01601 0.00922 -0.0493 0.0075 1.0000
9.250 1.2844 0.01649 0.00975 -0.0467 0.0070 1.0000
9.500 1.2994 0.01688 0.01018 -0.0446 0.0068 1.0000
9.750 1.3145 0.01729 0.01062 -0.0425 0.0064 1.0000
10.000 1.3287 0.01776 0.01112 -0.0404 0.0062 1.0000
10.250 1.3424 0.01828 0.01167 -0.0382 0.0059 1.0000
10.500 1.3556 0.01882 0.01227 -0.0360 0.0057 1.0000
10.750 1.3680 0.01943 0.01291 -0.0338 0.0054 1.0000
11.000 1.3787 0.02014 0.01367 -0.0314 0.0052 1.0000
11.250 1.3872 0.02100 0.01460 -0.0289 0.0051 1.0000
11.500 1.3967 0.02183 0.01549 -0.0265 0.0049 1.0000
11.750 1.4064 0.02268 0.01641 -0.0244 0.0048 1.0000
12.000 1.4144 0.02367 0.01747 -0.0222 0.0047 1.0000
12.250 1.4230 0.02466 0.01853 -0.0202 0.0045 1.0000
12.500 1.4296 0.02585 0.01979 -0.0182 0.0044 1.0000
12.750 1.4360 0.02711 0.02113 -0.0163 0.0043 1.0000
13.000 1.4423 0.02846 0.02255 -0.0147 0.0041 1.0000
13.250 1.4478 0.02995 0.02410 -0.0133 0.0040 1.0000
13.500 1.4515 0.03170 0.02593 -0.0120 0.0039 1.0000
13.750 1.4567 0.03340 0.02770 -0.0110 0.0038 1.0000
14.000 1.4583 0.03556 0.02994 -0.0101 0.0037 1.0000
14.250 1.4599 0.03785 0.03231 -0.0095 0.0036 1.0000
14.500 1.4566 0.04085 0.03542 -0.0091 0.0036 1.0000
14.750 1.4525 0.04416 0.03883 -0.0092 0.0035 1.0000
15.000 1.4458 0.04803 0.04280 -0.0097 0.0034 1.0000
15.250 1.4393 0.05206 0.04694 -0.0105 0.0034 1.0000
15.500 1.4316 0.05634 0.05135 -0.0114 0.0034 1.0000
15.750 1.4160 0.06181 0.05694 -0.0128 0.0033 1.0000
16.000 1.4062 0.06646 0.06169 -0.0139 0.0033 1.0000
16.250 1.3988 0.07070 0.06603 -0.0149 0.0033 1.0000
16.500 1.3828 0.07621 0.07165 -0.0163 0.0033 1.0000
16.750 1.3736 0.08077 0.07630 -0.0175 0.0033 1.0000
17.000 1.3647 0.08529 0.08092 -0.0186 0.0032 1.0000
17.250 1.3540 0.09018 0.08590 -0.0200 0.0032 1.0000
17.500 1.3437 0.09510 0.09090 -0.0214 0.0032 1.0000
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