USA 26 AIRFOIL (usa26-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 26 AIRFOIL (usa26-il) Reynolds number: 500,000 Max Cl/Cd: 92.9 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa26-il-500000-n5.txt Download as CSV file: xf-usa26-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 26 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.3791 0.12106 0.11870 -0.0265 1.0000 0.0106
-11.250 -0.3839 0.11632 0.11397 -0.0273 1.0000 0.0110
-11.000 -0.3910 0.11130 0.10898 -0.0281 1.0000 0.0114
-10.750 -0.3883 0.10912 0.10682 -0.0281 1.0000 0.0116
-10.500 -0.3846 0.10727 0.10499 -0.0278 1.0000 0.0119
-10.250 -0.3846 0.10464 0.10237 -0.0277 1.0000 0.0120
-10.000 -0.3853 0.10202 0.09977 -0.0276 1.0000 0.0124
-9.750 -0.3742 0.09838 0.09614 -0.0309 0.9989 0.0131
-9.500 -0.3689 0.09298 0.09074 -0.0353 0.9971 0.0138
-9.250 -0.3764 0.08442 0.08219 -0.0422 0.9944 0.0150
-9.000 -0.3659 0.08116 0.07893 -0.0455 0.9916 0.0152
-8.750 -0.3549 0.07781 0.07559 -0.0495 0.9886 0.0155
-8.500 -0.3418 0.07415 0.07193 -0.0548 0.9856 0.0159
-8.250 -0.3349 0.07028 0.06805 -0.0592 0.9795 0.0165
-7.750 -0.3352 0.04780 0.04526 -0.0800 0.9651 0.0196
-7.500 -0.3094 0.04483 0.04220 -0.0833 0.9630 0.0201
-7.250 -0.2944 0.04142 0.03867 -0.0843 0.9566 0.0206
-7.000 -0.2812 0.03249 0.02931 -0.0877 0.9512 0.0222
-6.750 -0.2808 0.02454 0.02050 -0.0861 0.9429 0.0241
-6.500 -0.2555 0.02253 0.01811 -0.0863 0.9369 0.0249
-6.250 -0.2376 0.01989 0.01510 -0.0853 0.9286 0.0256
-6.000 -0.2105 0.01873 0.01378 -0.0855 0.9209 0.0260
-5.750 -0.1853 0.01797 0.01289 -0.0852 0.9126 0.0265
-5.500 -0.1549 0.01731 0.01211 -0.0859 0.9048 0.0271
-5.250 -0.1274 0.01664 0.01129 -0.0859 0.8934 0.0278
-5.000 -0.0983 0.01577 0.01025 -0.0862 0.8820 0.0283
-4.750 -0.0690 0.01488 0.00916 -0.0866 0.8695 0.0287
-4.500 -0.0398 0.01413 0.00821 -0.0868 0.8537 0.0292
-4.250 -0.0116 0.01349 0.00738 -0.0868 0.8354 0.0295
-4.000 0.0152 0.01294 0.00665 -0.0865 0.8157 0.0300
-3.750 0.0406 0.01249 0.00603 -0.0858 0.7959 0.0303
-3.500 0.0652 0.01218 0.00558 -0.0850 0.7740 0.0309
-3.250 0.0889 0.01190 0.00513 -0.0840 0.7486 0.0314
-3.000 0.1112 0.01165 0.00470 -0.0826 0.7170 0.0317
-2.750 0.1316 0.01151 0.00434 -0.0809 0.6746 0.0320
-2.500 0.1510 0.01149 0.00407 -0.0790 0.6251 0.0323
-2.250 0.1704 0.01144 0.00379 -0.0772 0.5805 0.0326
-2.000 0.1912 0.01119 0.00337 -0.0756 0.5527 0.0333
-1.750 0.2136 0.01102 0.00309 -0.0744 0.5326 0.0340
-1.500 0.2369 0.01091 0.00290 -0.0734 0.5171 0.0347
-1.250 0.2606 0.01082 0.00274 -0.0725 0.5040 0.0354
-1.000 0.2847 0.01077 0.00262 -0.0716 0.4922 0.0362
-0.750 0.3090 0.01071 0.00251 -0.0708 0.4804 0.0371
-0.500 0.3334 0.01066 0.00241 -0.0701 0.4692 0.0383
-0.250 0.3579 0.01065 0.00235 -0.0693 0.4585 0.0399
0.000 0.3822 0.01063 0.00227 -0.0685 0.4480 0.0413
0.250 0.4070 0.01060 0.00220 -0.0678 0.4382 0.0431
0.500 0.4312 0.01057 0.00214 -0.0670 0.4290 0.0478
0.750 0.4558 0.01057 0.00211 -0.0662 0.4196 0.0541
1.000 0.4801 0.01055 0.00209 -0.0654 0.4112 0.0673
1.250 0.5022 0.01037 0.00211 -0.0643 0.4028 0.1548
1.500 0.5256 0.01029 0.00218 -0.0634 0.3954 0.2149
1.750 0.5493 0.01029 0.00225 -0.0626 0.3879 0.2562
2.000 0.5735 0.01030 0.00233 -0.0618 0.3812 0.2906
2.250 0.5967 0.01028 0.00242 -0.0609 0.3740 0.3406
2.500 0.6186 0.01015 0.00249 -0.0598 0.3675 0.4209
2.750 0.6916 0.00935 0.00298 -0.0699 0.3553 0.9710
3.000 0.7289 0.00953 0.00312 -0.0720 0.3471 0.9786
3.500 0.8044 0.00990 0.00340 -0.0765 0.3353 0.9901
3.750 0.8497 0.01009 0.00356 -0.0805 0.3296 0.9959
4.000 0.8910 0.01027 0.00370 -0.0837 0.3243 0.9994
4.250 0.9177 0.01038 0.00381 -0.0836 0.3204 1.0000
4.500 0.9392 0.01051 0.00394 -0.0823 0.3157 1.0000
4.750 0.9599 0.01067 0.00407 -0.0808 0.3109 1.0000
5.000 0.9812 0.01079 0.00422 -0.0795 0.3055 1.0000
5.250 1.0019 0.01094 0.00436 -0.0781 0.2984 1.0000
5.500 1.0223 0.01111 0.00451 -0.0766 0.2916 1.0000
5.750 1.0427 0.01127 0.00466 -0.0752 0.2829 1.0000
6.000 1.0628 0.01144 0.00484 -0.0737 0.2744 1.0000
6.250 1.0819 0.01166 0.00501 -0.0720 0.2591 1.0000
6.500 1.0991 0.01199 0.00521 -0.0700 0.2322 1.0000
6.750 1.1099 0.01273 0.00563 -0.0670 0.1784 1.0000
7.000 1.1220 0.01340 0.00613 -0.0642 0.1489 1.0000
7.250 1.1362 0.01393 0.00656 -0.0617 0.1289 1.0000
7.500 1.1492 0.01453 0.00701 -0.0591 0.1049 1.0000
7.750 1.1515 0.01573 0.00786 -0.0547 0.0492 1.0000
8.000 1.1553 0.01668 0.00866 -0.0504 0.0180 1.0000
8.250 1.1677 0.01710 0.00912 -0.0475 0.0151 1.0000
8.500 1.1813 0.01748 0.00956 -0.0449 0.0141 1.0000
8.750 1.1944 0.01791 0.01005 -0.0424 0.0132 1.0000
9.000 1.2060 0.01845 0.01065 -0.0396 0.0119 1.0000
9.250 1.2168 0.01906 0.01134 -0.0368 0.0111 1.0000
9.500 1.2283 0.01966 0.01201 -0.0341 0.0107 1.0000
9.750 1.2407 0.02023 0.01266 -0.0318 0.0103 1.0000
10.000 1.2521 0.02087 0.01337 -0.0293 0.0099 1.0000
10.250 1.2628 0.02157 0.01415 -0.0269 0.0093 1.0000
10.500 1.2722 0.02237 0.01501 -0.0244 0.0090 1.0000
10.750 1.2819 0.02317 0.01587 -0.0221 0.0084 1.0000
11.000 1.2877 0.02424 0.01702 -0.0194 0.0081 1.0000
11.250 1.2888 0.02566 0.01853 -0.0164 0.0078 1.0000
11.500 1.2924 0.02699 0.01996 -0.0139 0.0077 1.0000
11.750 1.2989 0.02820 0.02127 -0.0120 0.0075 1.0000
12.000 1.3025 0.02970 0.02287 -0.0100 0.0073 1.0000
12.250 1.3047 0.03141 0.02467 -0.0082 0.0071 1.0000
12.500 1.3057 0.03333 0.02670 -0.0066 0.0070 1.0000
12.750 1.3047 0.03560 0.02907 -0.0053 0.0068 1.0000
13.000 1.3049 0.03791 0.03149 -0.0043 0.0067 1.0000
13.250 1.3029 0.04063 0.03431 -0.0037 0.0065 1.0000
13.500 1.3016 0.04345 0.03723 -0.0035 0.0064 1.0000
13.750 1.2964 0.04692 0.04082 -0.0037 0.0063 1.0000
14.000 1.2928 0.05042 0.04442 -0.0041 0.0062 1.0000
14.250 1.2897 0.05398 0.04809 -0.0048 0.0060 1.0000
14.500 1.2824 0.05816 0.05239 -0.0057 0.0059 1.0000
14.750 1.2768 0.06218 0.05650 -0.0066 0.0059 1.0000
15.000 1.2709 0.06625 0.06065 -0.0076 0.0058 1.0000
15.250 1.2600 0.07099 0.06549 -0.0087 0.0056 1.0000
15.500 1.2518 0.07541 0.07001 -0.0098 0.0056 1.0000
15.750 1.2401 0.08033 0.07501 -0.0111 0.0054 1.0000
16.000 1.2334 0.08460 0.07936 -0.0122 0.0055 1.0000
16.250 1.2266 0.08889 0.08373 -0.0132 0.0055 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 26 AIRFOIL (usa26-il)