USA 25 AIRFOIL (usa25-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 25 AIRFOIL (usa25-il) Reynolds number: 500,000 Max Cl/Cd: 97.33 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa25-il-500000-n5.txt Download as CSV file: xf-usa25-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.2818 0.08842 0.08614 -0.0460 0.9869 0.0144
-8.500 -0.2722 0.08405 0.08177 -0.0504 0.9844 0.0153
-8.000 -0.2692 0.07241 0.07014 -0.0623 0.9737 0.0167
-7.750 -0.2558 0.07038 0.06810 -0.0640 0.9676 0.0169
-7.500 -0.2358 0.06738 0.06507 -0.0681 0.9638 0.0173
-7.250 -0.2209 0.06434 0.06200 -0.0709 0.9577 0.0178
-7.000 -0.2029 0.06065 0.05826 -0.0748 0.9512 0.0188
-6.750 -0.1888 0.05508 0.05259 -0.0794 0.9429 0.0198
-6.250 -0.1476 0.04494 0.04217 -0.0875 0.9225 0.0219
-6.000 -0.1175 0.04205 0.03916 -0.0909 0.9069 0.0224
-5.500 -0.0921 0.02022 0.01584 -0.0954 0.8654 0.0251
-5.250 -0.0651 0.01797 0.01308 -0.0959 0.8407 0.0254
-5.000 -0.0405 0.01674 0.01145 -0.0954 0.8151 0.0257
-4.750 -0.0177 0.01588 0.01027 -0.0945 0.7898 0.0259
-4.500 0.0043 0.01521 0.00930 -0.0932 0.7611 0.0261
-4.250 0.0247 0.01474 0.00850 -0.0916 0.7184 0.0263
-4.000 0.0416 0.01462 0.00797 -0.0891 0.6486 0.0265
-3.750 0.0592 0.01466 0.00762 -0.0869 0.5805 0.0267
-3.500 0.0794 0.01446 0.00710 -0.0853 0.5325 0.0269
-3.250 0.1012 0.01381 0.00617 -0.0841 0.5043 0.0271
-3.000 0.1244 0.01329 0.00547 -0.0831 0.4861 0.0273
-2.750 0.1482 0.01282 0.00485 -0.0822 0.4732 0.0278
-2.500 0.1726 0.01250 0.00443 -0.0815 0.4631 0.0282
-2.250 0.1972 0.01222 0.00408 -0.0808 0.4556 0.0286
-2.000 0.2220 0.01199 0.00380 -0.0800 0.4489 0.0290
-1.750 0.2467 0.01181 0.00356 -0.0793 0.4432 0.0294
-1.500 0.2718 0.01164 0.00334 -0.0787 0.4380 0.0297
-1.250 0.2968 0.01151 0.00317 -0.0780 0.4332 0.0301
-1.000 0.3215 0.01140 0.00301 -0.0773 0.4292 0.0305
-0.750 0.3467 0.01130 0.00288 -0.0766 0.4260 0.0309
-0.500 0.3720 0.01120 0.00276 -0.0760 0.4228 0.0317
-0.250 0.3971 0.01113 0.00266 -0.0753 0.4196 0.0324
0.000 0.4220 0.01109 0.00258 -0.0747 0.4164 0.0329
0.250 0.4468 0.01107 0.00252 -0.0740 0.4132 0.0333
0.500 0.4721 0.01104 0.00246 -0.0733 0.4107 0.0339
1.000 0.5226 0.01098 0.00239 -0.0721 0.4057 0.0369
1.250 0.5474 0.01096 0.00238 -0.0714 0.4021 0.0411
1.500 0.5714 0.01091 0.00241 -0.0706 0.3989 0.0725
1.750 0.5963 0.01095 0.00246 -0.0700 0.3959 0.0818
2.250 0.6468 0.01093 0.00257 -0.0689 0.3898 0.1076
2.500 0.6716 0.01095 0.00266 -0.0683 0.3864 0.1294
2.750 0.6963 0.01100 0.00274 -0.0676 0.3831 0.1449
3.000 0.7212 0.01107 0.00282 -0.0670 0.3802 0.1566
3.250 0.7467 0.01109 0.00290 -0.0665 0.3772 0.1680
3.500 0.7718 0.01112 0.00300 -0.0660 0.3739 0.1842
3.750 0.7964 0.01115 0.00310 -0.0653 0.3699 0.2076
4.000 0.8205 0.01122 0.00321 -0.0646 0.3663 0.2295
4.250 0.8454 0.01124 0.00333 -0.0640 0.3635 0.2556
4.500 0.8685 0.01111 0.00345 -0.0632 0.3605 0.3682
4.750 0.8990 0.01007 0.00376 -0.0639 0.3567 0.9527
5.000 0.9571 0.01032 0.00401 -0.0708 0.3499 0.9799
5.250 1.0041 0.01050 0.00416 -0.0752 0.3362 0.9937
5.500 1.0395 0.01068 0.00429 -0.0771 0.3168 0.9989
5.750 1.0631 0.01094 0.00442 -0.0765 0.2882 1.0000
6.000 1.0781 0.01136 0.00464 -0.0741 0.2525 1.0000
6.250 1.0905 0.01202 0.00503 -0.0713 0.2077 1.0000
6.500 1.1043 0.01264 0.00546 -0.0688 0.1762 1.0000
6.750 1.1202 0.01312 0.00584 -0.0667 0.1552 1.0000
7.000 1.1359 0.01363 0.00623 -0.0645 0.1340 1.0000
7.250 1.1388 0.01494 0.00709 -0.0603 0.0622 1.0000
7.750 1.1596 0.01644 0.00836 -0.0541 0.0159 1.0000
8.000 1.1738 0.01685 0.00883 -0.0515 0.0138 1.0000
8.250 1.1886 0.01722 0.00926 -0.0492 0.0132 1.0000
8.500 1.2026 0.01764 0.00974 -0.0467 0.0123 1.0000
8.750 1.2158 0.01814 0.01030 -0.0442 0.0113 1.0000
9.000 1.2283 0.01869 0.01092 -0.0417 0.0107 1.0000
9.250 1.2384 0.01939 0.01171 -0.0388 0.0099 1.0000
9.500 1.2515 0.01995 0.01234 -0.0365 0.0096 1.0000
9.750 1.2635 0.02058 0.01305 -0.0341 0.0093 1.0000
10.000 1.2745 0.02127 0.01381 -0.0317 0.0090 1.0000
10.250 1.2844 0.02205 0.01466 -0.0292 0.0085 1.0000
10.500 1.2938 0.02288 0.01556 -0.0268 0.0083 1.0000
10.750 1.3013 0.02383 0.01658 -0.0243 0.0078 1.0000
11.000 1.3075 0.02491 0.01774 -0.0218 0.0075 1.0000
11.250 1.3107 0.02623 0.01914 -0.0191 0.0074 1.0000
11.500 1.3104 0.02787 0.02088 -0.0164 0.0071 1.0000
11.750 1.3151 0.02928 0.02237 -0.0144 0.0069 1.0000
12.000 1.3165 0.03103 0.02423 -0.0125 0.0068 1.0000
12.250 1.3187 0.03285 0.02615 -0.0109 0.0066 1.0000
12.500 1.3176 0.03512 0.02852 -0.0095 0.0065 1.0000
12.750 1.3169 0.03752 0.03103 -0.0085 0.0063 1.0000
13.000 1.3145 0.04028 0.03390 -0.0079 0.0062 1.0000
13.250 1.3106 0.04343 0.03717 -0.0077 0.0061 1.0000
13.500 1.3062 0.04688 0.04073 -0.0079 0.0060 1.0000
13.750 1.3024 0.05042 0.04438 -0.0083 0.0059 1.0000
14.000 1.2983 0.05412 0.04818 -0.0090 0.0058 1.0000
14.250 1.2937 0.05795 0.05211 -0.0097 0.0057 1.0000
14.500 1.2870 0.06207 0.05633 -0.0105 0.0057 1.0000
14.750 1.2811 0.06608 0.06043 -0.0113 0.0056 1.0000
15.000 1.2779 0.06977 0.06421 -0.0122 0.0054 1.0000
15.250 1.2715 0.07386 0.06838 -0.0131 0.0054 1.0000
15.500 1.2648 0.07797 0.07258 -0.0139 0.0054 1.0000
15.750 1.2604 0.08188 0.07656 -0.0149 0.0053 1.0000
16.000 1.2549 0.08596 0.08071 -0.0159 0.0052 1.0000
16.250 1.2489 0.09012 0.08494 -0.0169 0.0051 1.0000
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Polar data table (+)
Polar graphs
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