USA 29 AIRFOIL (usa29-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: USA 29 AIRFOIL (usa29-il) Reynolds number: 200,000 Max Cl/Cd: 74.57 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa29-il-200000.txt Download as CSV file: xf-usa29-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: USA 29 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.1607 0.10364 0.10042 -0.0604 0.9797 0.0760
-10.000 -0.3173 0.10989 0.10663 -0.0507 0.9920 0.0709
-9.750 -0.2821 0.10757 0.10426 -0.0490 0.9900 0.0734
-9.500 -0.2634 0.10422 0.10088 -0.0524 0.9866 0.0773
-9.250 -0.2840 0.09892 0.09564 -0.0679 0.9779 0.0821
-9.000 -0.2569 0.09417 0.09085 -0.0656 0.9764 0.0838
-8.750 -0.2261 0.09188 0.08852 -0.0660 0.9738 0.0867
-8.000 -0.1999 0.07832 0.07496 -0.0844 0.9563 0.0979
-7.500 -0.1652 0.07274 0.06932 -0.0887 0.9436 0.1049
-7.250 -0.2275 0.04757 0.04365 -0.1093 0.9253 0.0839
-7.000 -0.2770 0.03318 0.02787 -0.1058 0.9105 0.0718
-6.750 -0.2627 0.03110 0.02556 -0.1041 0.9019 0.0727
-6.500 -0.2314 0.02944 0.02374 -0.1054 0.8976 0.0745
-6.250 -0.2009 0.02690 0.02078 -0.1065 0.8944 0.0753
-6.000 -0.1851 0.02533 0.01891 -0.1042 0.8848 0.0759
-5.750 -0.1526 0.02370 0.01699 -0.1052 0.8805 0.0775
-5.500 -0.1248 0.02248 0.01550 -0.1050 0.8740 0.0794
-5.250 -0.0968 0.02127 0.01403 -0.1048 0.8665 0.0805
-5.000 -0.0592 0.02007 0.01255 -0.1064 0.8625 0.0818
-4.750 -0.0375 0.01904 0.01139 -0.1049 0.8517 0.0830
-4.500 -0.0003 0.01795 0.01025 -0.1067 0.8467 0.0860
-4.250 0.0227 0.01745 0.00971 -0.1055 0.8351 0.0889
-4.000 0.0517 0.01685 0.00900 -0.1054 0.8257 0.0918
-3.750 0.0832 0.01628 0.00828 -0.1057 0.8168 0.0949
-3.500 0.1073 0.01561 0.00763 -0.1048 0.8052 0.0989
-3.250 0.1351 0.01516 0.00713 -0.1046 0.7951 0.1047
-3.000 0.1628 0.01467 0.00659 -0.1043 0.7848 0.1136
-2.750 0.1859 0.01422 0.00618 -0.1032 0.7729 0.1291
-2.500 0.2114 0.01371 0.00585 -0.1027 0.7624 0.1714
-2.250 0.2393 0.01346 0.00564 -0.1025 0.7527 0.2182
-2.000 0.2620 0.01341 0.00563 -0.1014 0.7411 0.2487
-1.750 0.2864 0.01349 0.00572 -0.1005 0.7311 0.2739
-1.500 0.3126 0.01363 0.00575 -0.0999 0.7219 0.2957
-1.250 0.3350 0.01383 0.00588 -0.0986 0.7113 0.3105
-1.000 0.3604 0.01402 0.00598 -0.0979 0.7027 0.3222
-0.750 0.3842 0.01419 0.00607 -0.0969 0.6935 0.3331
-0.500 0.4083 0.01439 0.00617 -0.0959 0.6849 0.3455
-0.250 0.4330 0.01459 0.00623 -0.0951 0.6767 0.3580
0.000 0.4573 0.01468 0.00631 -0.0944 0.6684 0.3671
0.250 0.4831 0.01476 0.00630 -0.0939 0.6609 0.3779
0.500 0.5071 0.01486 0.00635 -0.0931 0.6532 0.3896
0.750 0.5328 0.01491 0.00633 -0.0926 0.6458 0.4007
1.000 0.5589 0.01489 0.00631 -0.0923 0.6386 0.4097
1.250 0.5846 0.01489 0.00628 -0.0919 0.6314 0.4192
1.500 0.6120 0.01494 0.00627 -0.0919 0.6251 0.4294
1.750 0.6364 0.01488 0.00627 -0.0913 0.6177 0.4382
2.000 0.6655 0.01492 0.00624 -0.0917 0.6118 0.4481
2.250 0.6888 0.01492 0.00635 -0.0909 0.6049 0.4568
2.500 0.7155 0.01493 0.00637 -0.0907 0.5986 0.4656
2.750 0.7422 0.01500 0.00645 -0.0906 0.5928 0.4742
3.000 0.7648 0.01501 0.00657 -0.0897 0.5862 0.4820
3.250 0.7921 0.01507 0.00664 -0.0896 0.5808 0.4916
3.500 0.8147 0.01515 0.00682 -0.0887 0.5748 0.5011
3.750 0.8377 0.01520 0.00697 -0.0878 0.5686 0.5136
4.000 0.8660 0.01520 0.00707 -0.0880 0.5637 0.5487
4.250 1.0187 0.01517 0.00802 -0.1154 0.5531 0.9865
4.500 1.0818 0.01541 0.00818 -0.1232 0.5471 0.9992
4.750 1.1035 0.01565 0.00853 -0.1221 0.5409 1.0000
5.000 1.1257 0.01587 0.00874 -0.1211 0.5354 1.0000
5.250 1.1514 0.01612 0.00893 -0.1207 0.5307 1.0000
5.500 1.1681 0.01640 0.00934 -0.1186 0.5248 1.0000
5.750 1.1900 0.01661 0.00955 -0.1175 0.5192 1.0000
6.000 1.2093 0.01678 0.00973 -0.1158 0.5118 1.0000
6.250 1.2264 0.01679 0.00974 -0.1136 0.5020 1.0000
6.500 1.2400 0.01690 0.00988 -0.1107 0.4920 1.0000
6.750 1.2602 0.01700 0.00991 -0.1091 0.4837 1.0000
7.000 1.2704 0.01713 0.01014 -0.1057 0.4737 1.0000
7.250 1.2844 0.01726 0.01029 -0.1029 0.4641 1.0000
7.500 1.2976 0.01740 0.01043 -0.1000 0.4542 1.0000
7.750 1.3054 0.01758 0.01070 -0.0961 0.4428 1.0000
8.000 1.3110 0.01776 0.01089 -0.0917 0.4293 1.0000
8.250 1.3141 0.01797 0.01112 -0.0868 0.4152 1.0000
8.500 1.3107 0.01820 0.01134 -0.0807 0.3992 1.0000
8.750 1.3044 0.01858 0.01165 -0.0744 0.3778 1.0000
9.000 1.2948 0.01926 0.01222 -0.0678 0.3457 1.0000
9.250 1.2785 0.02051 0.01319 -0.0608 0.2934 1.0000
9.500 1.2594 0.02235 0.01467 -0.0542 0.2372 1.0000
9.750 1.2315 0.02512 0.01696 -0.0474 0.1680 1.0000
10.000 1.2020 0.02846 0.01986 -0.0414 0.0982 1.0000
10.250 1.1814 0.03159 0.02274 -0.0370 0.0587 1.0000
10.500 1.1745 0.03398 0.02511 -0.0342 0.0500 1.0000
10.750 1.1720 0.03617 0.02739 -0.0319 0.0457 1.0000
11.000 1.1713 0.03833 0.02966 -0.0300 0.0435 1.0000
11.250 1.1672 0.04091 0.03235 -0.0282 0.0415 1.0000
11.500 1.1604 0.04391 0.03546 -0.0266 0.0400 1.0000
11.750 1.1535 0.04705 0.03872 -0.0253 0.0386 1.0000
12.000 1.1515 0.04982 0.04160 -0.0243 0.0379 1.0000
12.250 1.1479 0.05283 0.04472 -0.0234 0.0366 1.0000
12.500 1.1440 0.05592 0.04790 -0.0226 0.0357 1.0000
12.750 1.1402 0.05903 0.05109 -0.0218 0.0351 1.0000
13.000 1.1374 0.06205 0.05417 -0.0211 0.0343 1.0000
13.250 1.1349 0.06500 0.05715 -0.0203 0.0336 1.0000
13.500 1.1344 0.06761 0.05976 -0.0193 0.0328 1.0000
13.750 1.1383 0.06948 0.06161 -0.0179 0.0321 1.0000
14.000 1.1499 0.07027 0.06234 -0.0158 0.0314 1.0000
14.250 1.1588 0.07197 0.06413 -0.0151 0.0307 1.0000
14.500 1.1690 0.07342 0.06566 -0.0141 0.0297 1.0000
14.750 1.1842 0.07417 0.06644 -0.0126 0.0292 1.0000
15.000 1.2010 0.07482 0.06713 -0.0111 0.0286 1.0000
15.250 1.2200 0.07533 0.06769 -0.0094 0.0280 1.0000
15.500 1.2396 0.07597 0.06842 -0.0078 0.0279 1.0000
15.750 1.2592 0.07685 0.06942 -0.0063 0.0277 1.0000
16.000 1.2751 0.07824 0.07092 -0.0051 0.0274 1.0000
16.250 1.2962 0.07929 0.07199 -0.0040 0.0267 1.0000
16.500 1.3112 0.08116 0.07395 -0.0030 0.0263 1.0000
16.750 1.3196 0.08373 0.07672 -0.0023 0.0264 1.0000
17.000 1.3084 0.08794 0.08126 -0.0020 0.0268 1.0000
17.250 1.2914 0.09310 0.08682 -0.0022 0.0278 1.0000
17.500 1.2845 0.09786 0.09189 -0.0021 0.0292 1.0000
17.750 1.2794 0.10258 0.09682 -0.0021 0.0302 1.0000
18.000 1.0433 0.10105 0.09591 0.0054 0.0287 1.0000
18.250 1.0266 0.10510 0.10018 0.0042 0.0291 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 29 AIRFOIL (usa29-il)