USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 100,000 Max Cl/Cd: 55.71 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa22-il-100000-n5.txt Download as CSV file: xf-usa22-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3364 0.09076 0.08625 -0.0217 1.0000 0.0306
-7.500 -0.3347 0.08744 0.08300 -0.0227 1.0000 0.0294
-7.250 -0.3338 0.08398 0.07962 -0.0244 1.0000 0.0284
-7.000 -0.3314 0.08010 0.07581 -0.0270 1.0000 0.0275
-6.750 -0.3292 0.07589 0.07167 -0.0300 1.0000 0.0266
-6.500 -0.3272 0.07082 0.06667 -0.0340 1.0000 0.0256
-6.000 -0.3156 0.05961 0.05545 -0.0418 1.0000 0.0245
-5.750 -0.2960 0.05811 0.05393 -0.0434 0.9954 0.0262
-5.500 -0.2586 0.05255 0.04821 -0.0516 0.9854 0.0281
-5.250 -0.2181 0.04373 0.03903 -0.0616 0.9747 0.0285
-5.000 -0.1776 0.03450 0.02907 -0.0699 0.9649 0.0291
-4.750 -0.1405 0.02855 0.02216 -0.0743 0.9540 0.0315
-4.500 -0.1055 0.02543 0.01839 -0.0766 0.9422 0.0351
-4.250 -0.0705 0.02343 0.01601 -0.0783 0.9301 0.0385
-4.000 -0.0348 0.02168 0.01367 -0.0796 0.9173 0.0450
-3.750 -0.0015 0.02020 0.01203 -0.0807 0.9032 0.0508
-3.500 0.0315 0.01917 0.01062 -0.0814 0.8876 0.0591
-3.250 0.0622 0.01817 0.00955 -0.0818 0.8705 0.0669
-3.000 0.0926 0.01739 0.00849 -0.0818 0.8524 0.0731
-2.750 0.1204 0.01661 0.00765 -0.0815 0.8319 0.0784
-2.500 0.1486 0.01610 0.00696 -0.0811 0.8112 0.0856
-2.250 0.1762 0.01562 0.00642 -0.0807 0.7899 0.0968
-2.000 0.2038 0.01521 0.00593 -0.0802 0.7689 0.1095
-1.750 0.2312 0.01482 0.00554 -0.0797 0.7488 0.1393
-1.500 0.2581 0.01453 0.00539 -0.0792 0.7285 0.2026
-1.250 0.2853 0.01454 0.00532 -0.0788 0.7090 0.2587
-1.000 0.3122 0.01457 0.00521 -0.0782 0.6905 0.2916
-0.750 0.3389 0.01458 0.00512 -0.0778 0.6719 0.3246
-0.500 0.3655 0.01454 0.00501 -0.0773 0.6540 0.3589
-0.250 0.3921 0.01443 0.00486 -0.0769 0.6370 0.3907
0.000 0.4185 0.01428 0.00471 -0.0765 0.6209 0.4230
0.250 0.4424 0.01372 0.00465 -0.0757 0.6058 0.5524
0.750 0.5053 0.01327 0.00451 -0.0762 0.5761 1.0000
1.000 0.5319 0.01347 0.00452 -0.0758 0.5628 1.0000
1.250 0.5584 0.01368 0.00456 -0.0754 0.5501 1.0000
1.500 0.5849 0.01390 0.00462 -0.0750 0.5380 1.0000
1.750 0.6112 0.01413 0.00469 -0.0746 0.5265 1.0000
2.000 0.6375 0.01436 0.00478 -0.0742 0.5150 1.0000
2.250 0.6637 0.01460 0.00492 -0.0739 0.5031 1.0000
2.500 0.6899 0.01484 0.00507 -0.0735 0.4920 1.0000
2.750 0.7160 0.01509 0.00523 -0.0731 0.4815 1.0000
3.000 0.7420 0.01536 0.00542 -0.0727 0.4713 1.0000
3.250 0.7681 0.01562 0.00564 -0.0724 0.4608 1.0000
3.500 0.7941 0.01589 0.00587 -0.0720 0.4513 1.0000
3.750 0.8200 0.01618 0.00613 -0.0717 0.4421 1.0000
4.000 0.8459 0.01647 0.00643 -0.0713 0.4325 1.0000
4.250 0.8716 0.01678 0.00671 -0.0709 0.4240 1.0000
4.500 0.8973 0.01708 0.00707 -0.0706 0.4146 1.0000
4.750 0.9229 0.01739 0.00742 -0.0702 0.4053 1.0000
5.000 0.9482 0.01773 0.00774 -0.0698 0.3965 1.0000
5.250 0.9735 0.01804 0.00818 -0.0694 0.3863 1.0000
5.500 0.9987 0.01839 0.00859 -0.0690 0.3772 1.0000
5.750 1.0236 0.01875 0.00900 -0.0685 0.3682 1.0000
6.000 1.0485 0.01911 0.00950 -0.0681 0.3581 1.0000
6.250 1.0731 0.01949 0.00999 -0.0675 0.3486 1.0000
6.500 1.0972 0.01987 0.01044 -0.0670 0.3382 1.0000
6.750 1.1199 0.02016 0.01083 -0.0662 0.3196 1.0000
7.000 1.1421 0.02050 0.01121 -0.0654 0.2992 1.0000
7.250 1.1626 0.02092 0.01159 -0.0645 0.2714 1.0000
7.500 1.1831 0.02150 0.01219 -0.0636 0.2467 1.0000
7.750 1.2024 0.02225 0.01287 -0.0625 0.2186 1.0000
8.000 1.2188 0.02331 0.01375 -0.0613 0.1802 1.0000
8.250 1.2323 0.02476 0.01491 -0.0598 0.1178 1.0000
8.500 1.2260 0.02829 0.01762 -0.0566 0.0303 1.0000
8.750 1.2372 0.02985 0.01931 -0.0548 0.0257 1.0000
9.000 1.2489 0.03126 0.02095 -0.0530 0.0239 1.0000
9.250 1.2577 0.03283 0.02276 -0.0510 0.0224 1.0000
9.500 1.2614 0.03459 0.02474 -0.0486 0.0211 1.0000
9.750 1.2604 0.03671 0.02707 -0.0462 0.0199 1.0000
10.000 1.2550 0.03933 0.02997 -0.0442 0.0189 1.0000
10.250 1.2469 0.04249 0.03337 -0.0430 0.0183 1.0000
10.500 1.2372 0.04620 0.03731 -0.0427 0.0179 1.0000
10.750 1.2303 0.05001 0.04128 -0.0430 0.0178 1.0000
11.000 1.2230 0.05415 0.04562 -0.0439 0.0176 1.0000
11.250 1.2152 0.05857 0.05022 -0.0450 0.0175 1.0000
11.500 1.2071 0.06313 0.05494 -0.0462 0.0174 1.0000
11.750 1.1996 0.06764 0.05960 -0.0473 0.0172 1.0000
12.000 1.1933 0.07193 0.06402 -0.0482 0.0171 1.0000
12.250 1.1889 0.07582 0.06803 -0.0488 0.0169 1.0000
12.500 1.1873 0.07917 0.07147 -0.0489 0.0167 1.0000
12.750 1.1881 0.08204 0.07443 -0.0486 0.0164 1.0000
13.000 1.1912 0.08460 0.07709 -0.0481 0.0159 1.0000
13.250 1.1945 0.08723 0.07980 -0.0476 0.0153 1.0000
13.500 1.1987 0.08972 0.08240 -0.0471 0.0147 1.0000
13.750 1.2034 0.09220 0.08496 -0.0465 0.0142 1.0000
14.000 1.2092 0.09460 0.08748 -0.0457 0.0139 1.0000
14.250 1.2131 0.09747 0.09049 -0.0453 0.0137 1.0000
14.500 1.2143 0.10093 0.09411 -0.0455 0.0136 1.0000
14.750 1.2136 0.10485 0.09820 -0.0462 0.0135 1.0000
15.000 1.2105 0.10931 0.10284 -0.0475 0.0135 1.0000
15.250 1.2057 0.11421 0.10794 -0.0493 0.0134 1.0000
15.500 1.1993 0.11955 0.11353 -0.0516 0.0134 1.0000
15.750 1.1914 0.12534 0.11952 -0.0545 0.0135 1.0000
16.000 1.1824 0.13160 0.12597 -0.0579 0.0135 1.0000
16.250 1.1724 0.13834 0.13290 -0.0618 0.0136 1.0000
16.500 1.1616 0.14559 0.14033 -0.0662 0.0137 1.0000
16.750 1.1500 0.15338 0.14830 -0.0712 0.0138 1.0000
17.000 1.1378 0.16178 0.15686 -0.0766 0.0139 1.0000
17.250 1.1250 0.17091 0.16608 -0.0826 0.0141 1.0000
17.500 1.1117 0.18090 0.17620 -0.0891 0.0143 1.0000
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Polar data table (+)
Polar graphs
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