USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 1,000,000 Max Cl/Cd: 112.55 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa22-il-1000000-n5.txt Download as CSV file: xf-usa22-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 22 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4234 0.12509 0.12348 -0.0006 1.0000 0.0033
-10.750 -0.4196 0.12098 0.11939 -0.0021 1.0000 0.0033
-9.250 -0.5851 0.02359 0.02092 -0.0765 0.9866 0.0036
-9.000 -0.5559 0.02079 0.01781 -0.0791 0.9763 0.0037
-8.750 -0.5260 0.01887 0.01564 -0.0809 0.9619 0.0038
-8.500 -0.4972 0.01738 0.01391 -0.0820 0.9445 0.0040
-8.250 -0.4709 0.01628 0.01258 -0.0821 0.9265 0.0041
-8.000 -0.4457 0.01538 0.01146 -0.0819 0.9103 0.0043
-7.750 -0.4204 0.01461 0.01049 -0.0816 0.8954 0.0045
-7.500 -0.3948 0.01394 0.00964 -0.0813 0.8797 0.0047
-7.250 -0.3689 0.01333 0.00885 -0.0809 0.8629 0.0048
-7.000 -0.3435 0.01242 0.00769 -0.0806 0.8444 0.0050
-6.750 -0.3177 0.01183 0.00690 -0.0802 0.8219 0.0053
-6.500 -0.2917 0.01144 0.00633 -0.0798 0.7916 0.0056
-6.250 -0.2656 0.01112 0.00581 -0.0794 0.7599 0.0059
-6.000 -0.2391 0.01082 0.00533 -0.0791 0.7344 0.0063
-5.750 -0.2121 0.01051 0.00488 -0.0788 0.7154 0.0067
-5.500 -0.1848 0.01025 0.00448 -0.0786 0.6986 0.0070
-5.250 -0.1575 0.00994 0.00403 -0.0784 0.6822 0.0075
-5.000 -0.1302 0.00965 0.00363 -0.0782 0.6657 0.0086
-4.750 -0.1027 0.00944 0.00333 -0.0780 0.6501 0.0096
-4.500 -0.0750 0.00927 0.00305 -0.0779 0.6355 0.0104
-4.250 -0.0473 0.00906 0.00274 -0.0777 0.6199 0.0123
-4.000 -0.0198 0.00890 0.00252 -0.0775 0.6031 0.0149
-3.750 0.0079 0.00877 0.00234 -0.0774 0.5857 0.0207
-3.500 0.0356 0.00869 0.00219 -0.0773 0.5690 0.0247
-3.250 0.0634 0.00863 0.00208 -0.0772 0.5517 0.0292
-3.000 0.0912 0.00861 0.00198 -0.0770 0.5337 0.0323
-2.750 0.1190 0.00862 0.00190 -0.0769 0.5165 0.0336
-2.500 0.1467 0.00854 0.00176 -0.0768 0.5011 0.0374
-2.250 0.1745 0.00852 0.00168 -0.0767 0.4864 0.0405
-2.000 0.2024 0.00851 0.00160 -0.0766 0.4735 0.0427
-1.750 0.2304 0.00849 0.00151 -0.0765 0.4625 0.0438
-1.500 0.2583 0.00846 0.00142 -0.0764 0.4522 0.0460
-1.250 0.2862 0.00842 0.00133 -0.0763 0.4426 0.0505
-1.000 0.3142 0.00840 0.00128 -0.0762 0.4341 0.0547
-0.750 0.3421 0.00841 0.00123 -0.0762 0.4250 0.0586
-0.500 0.3699 0.00837 0.00119 -0.0761 0.4144 0.0722
-0.250 0.3973 0.00820 0.00117 -0.0761 0.4044 0.1391
0.000 0.4250 0.00820 0.00117 -0.0760 0.3958 0.1598
0.250 0.4530 0.00815 0.00117 -0.0760 0.3900 0.1841
0.500 0.4807 0.00812 0.00121 -0.0759 0.3832 0.2231
0.750 0.5088 0.00813 0.00124 -0.0759 0.3775 0.2442
1.000 0.5367 0.00814 0.00128 -0.0759 0.3718 0.2631
1.250 0.5646 0.00819 0.00133 -0.0758 0.3665 0.2774
1.750 0.6206 0.00826 0.00142 -0.0758 0.3555 0.3027
2.000 0.6484 0.00831 0.00148 -0.0758 0.3500 0.3123
2.250 0.6763 0.00837 0.00154 -0.0757 0.3430 0.3190
2.500 0.7040 0.00844 0.00160 -0.0757 0.3357 0.3278
2.750 0.7317 0.00851 0.00168 -0.0757 0.3275 0.3389
3.000 0.7593 0.00856 0.00176 -0.0756 0.3204 0.3584
3.250 0.7868 0.00861 0.00185 -0.0756 0.3121 0.3869
3.500 0.8090 0.00771 0.00204 -0.0748 0.3047 0.8347
4.000 0.8682 0.00773 0.00226 -0.0755 0.2851 1.0000
4.250 0.8948 0.00795 0.00241 -0.0753 0.2696 1.0000
4.500 0.9208 0.00825 0.00260 -0.0751 0.2473 1.0000
4.750 0.9471 0.00851 0.00278 -0.0749 0.2315 1.0000
5.000 0.9728 0.00885 0.00300 -0.0746 0.2115 1.0000
5.250 0.9983 0.00920 0.00325 -0.0743 0.1914 1.0000
5.500 1.0238 0.00955 0.00351 -0.0740 0.1727 1.0000
5.750 1.0489 0.00996 0.00380 -0.0737 0.1522 1.0000
6.000 1.0682 0.01114 0.00454 -0.0727 0.0788 1.0000
6.250 1.0880 0.01228 0.00541 -0.0717 0.0158 1.0000
6.500 1.1132 0.01262 0.00576 -0.0713 0.0113 1.0000
6.750 1.1387 0.01292 0.00609 -0.0710 0.0100 1.0000
7.000 1.1638 0.01325 0.00645 -0.0707 0.0088 1.0000
7.250 1.1882 0.01367 0.00690 -0.0702 0.0075 1.0000
7.500 1.2127 0.01407 0.00733 -0.0698 0.0067 1.0000
7.750 1.2371 0.01444 0.00773 -0.0694 0.0060 1.0000
8.000 1.2610 0.01486 0.00817 -0.0689 0.0055 1.0000
8.250 1.2840 0.01538 0.00872 -0.0683 0.0049 1.0000
8.500 1.3066 0.01592 0.00932 -0.0677 0.0046 1.0000
8.750 1.3293 0.01643 0.00988 -0.0671 0.0043 1.0000
9.000 1.3515 0.01696 0.01048 -0.0664 0.0040 1.0000
9.250 1.3733 0.01749 0.01105 -0.0657 0.0037 1.0000
9.500 1.3948 0.01803 0.01162 -0.0650 0.0035 1.0000
9.750 1.4147 0.01870 0.01234 -0.0641 0.0032 1.0000
10.000 1.4320 0.01959 0.01332 -0.0628 0.0030 1.0000
10.250 1.4504 0.02032 0.01411 -0.0617 0.0029 1.0000
10.500 1.4672 0.02112 0.01500 -0.0604 0.0028 1.0000
10.750 1.4824 0.02200 0.01596 -0.0589 0.0027 1.0000
11.000 1.4957 0.02292 0.01697 -0.0572 0.0026 1.0000
11.250 1.5046 0.02392 0.01807 -0.0549 0.0025 1.0000
11.500 1.5117 0.02503 0.01928 -0.0525 0.0024 1.0000
11.750 1.5180 0.02627 0.02060 -0.0504 0.0023 1.0000
12.000 1.5229 0.02774 0.02216 -0.0486 0.0023 1.0000
12.250 1.5270 0.02942 0.02393 -0.0471 0.0022 1.0000
12.500 1.5294 0.03143 0.02605 -0.0459 0.0022 1.0000
12.750 1.5315 0.03369 0.02840 -0.0452 0.0021 1.0000
13.000 1.5322 0.03635 0.03116 -0.0450 0.0021 1.0000
13.250 1.5296 0.03962 0.03455 -0.0453 0.0020 1.0000
13.500 1.5263 0.04321 0.03825 -0.0459 0.0020 1.0000
13.750 1.5190 0.04745 0.04261 -0.0468 0.0020 1.0000
14.000 1.5094 0.05215 0.04744 -0.0480 0.0019 1.0000
14.250 1.4962 0.05743 0.05285 -0.0494 0.0019 1.0000
14.500 1.4829 0.06277 0.05832 -0.0508 0.0019 1.0000
14.750 1.4677 0.06841 0.06408 -0.0523 0.0019 1.0000
15.000 1.4515 0.07426 0.07005 -0.0540 0.0018 1.0000
15.250 1.4338 0.08044 0.07635 -0.0557 0.0018 1.0000
15.500 1.4212 0.08594 0.08195 -0.0573 0.0018 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 22 AIRFOIL (usa22-il)