USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 100,000 Max Cl/Cd: 53.78 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa22-il-100000.txt Download as CSV file: xf-usa22-il-100000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3690 0.10763 0.10292 -0.0115 1.0000 0.0757
-8.500 -0.3739 0.10603 0.10141 -0.0161 1.0000 0.0778
-8.250 -0.3822 0.10466 0.10015 -0.0215 1.0000 0.0784
-8.000 -0.3678 0.09848 0.09399 -0.0190 1.0000 0.0800
-7.750 -0.3520 0.09456 0.09007 -0.0169 1.0000 0.0831
-7.500 -0.3469 0.09166 0.08722 -0.0176 1.0000 0.0864
-7.250 -0.3474 0.08909 0.08471 -0.0202 1.0000 0.0899
-7.000 -0.3490 0.08693 0.08262 -0.0354 1.0000 0.0925
-6.750 -0.3403 0.08202 0.07779 -0.0273 1.0000 0.0945
-6.500 -0.3308 0.07916 0.07497 -0.0251 1.0000 0.0981
-6.250 -0.3255 0.07633 0.07219 -0.0290 1.0000 0.1041
-6.000 -0.2909 0.06328 0.05970 -0.0213 1.0000 0.1118
-5.750 -0.2996 0.06148 0.05798 -0.0203 1.0000 0.1157
-5.500 -0.3076 0.06704 0.06293 -0.0367 1.0000 0.1209
-5.250 -0.3056 0.06481 0.06085 -0.0305 1.0000 0.1248
-5.000 -0.2963 0.06200 0.05792 -0.0355 1.0000 0.1353
-4.750 -0.2953 0.05977 0.05582 -0.0312 1.0000 0.1382
-4.500 -0.2841 0.05733 0.05324 -0.0348 1.0000 0.1499
-4.250 -0.2713 0.05463 0.05064 -0.0331 0.9975 0.1546
-4.000 -0.2156 0.04995 0.04565 -0.0439 0.9878 0.1795
-3.750 -0.1407 0.03319 0.02733 -0.0599 0.9835 0.0947
-3.500 -0.0909 0.02770 0.02102 -0.0647 0.9747 0.0867
-3.250 -0.0394 0.02453 0.01690 -0.0690 0.9651 0.0920
-3.000 0.0118 0.02204 0.01416 -0.0736 0.9565 0.0997
-2.750 0.0604 0.02023 0.01197 -0.0771 0.9445 0.1089
-2.500 0.1070 0.01899 0.01062 -0.0802 0.9313 0.1226
-2.250 0.1505 0.01784 0.00941 -0.0825 0.9167 0.1363
-2.000 0.1901 0.01690 0.00854 -0.0840 0.9002 0.1601
-1.750 0.2234 0.01599 0.00788 -0.0841 0.8806 0.2080
-1.500 0.2527 0.01579 0.00789 -0.0835 0.8582 0.3161
-1.250 0.2800 0.01549 0.00749 -0.0825 0.8355 0.3653
-1.000 0.3057 0.01503 0.00705 -0.0813 0.8129 0.4085
-0.750 0.3306 0.01447 0.00664 -0.0800 0.7899 0.4588
-0.500 0.3676 0.01275 0.00628 -0.0804 0.7686 1.0000
-0.250 0.3944 0.01296 0.00609 -0.0796 0.7463 1.0000
0.000 0.4208 0.01318 0.00597 -0.0787 0.7262 1.0000
0.250 0.4471 0.01344 0.00592 -0.0779 0.7079 1.0000
0.500 0.4730 0.01373 0.00596 -0.0773 0.6893 1.0000
0.750 0.4990 0.01402 0.00603 -0.0766 0.6725 1.0000
1.000 0.5251 0.01431 0.00612 -0.0760 0.6570 1.0000
1.250 0.5512 0.01460 0.00624 -0.0754 0.6423 1.0000
1.500 0.5772 0.01489 0.00637 -0.0749 0.6283 1.0000
1.750 0.6033 0.01518 0.00652 -0.0743 0.6148 1.0000
2.000 0.6294 0.01546 0.00667 -0.0738 0.6019 1.0000
2.250 0.6556 0.01575 0.00683 -0.0732 0.5895 1.0000
2.500 0.6817 0.01603 0.00700 -0.0727 0.5773 1.0000
2.750 0.7075 0.01633 0.00724 -0.0722 0.5646 1.0000
3.000 0.7334 0.01666 0.00752 -0.0717 0.5524 1.0000
3.250 0.7593 0.01699 0.00781 -0.0713 0.5408 1.0000
3.500 0.7854 0.01732 0.00807 -0.0708 0.5296 1.0000
3.750 0.8116 0.01765 0.00832 -0.0703 0.5188 1.0000
4.000 0.8370 0.01802 0.00871 -0.0698 0.5067 1.0000
4.250 0.8625 0.01842 0.00914 -0.0693 0.4950 1.0000
4.500 0.8882 0.01883 0.00954 -0.0689 0.4837 1.0000
4.750 0.9140 0.01923 0.00990 -0.0684 0.4725 1.0000
5.000 0.9396 0.01963 0.01030 -0.0679 0.4609 1.0000
5.250 0.9644 0.02009 0.01082 -0.0673 0.4483 1.0000
5.500 0.9892 0.02060 0.01140 -0.0668 0.4361 1.0000
5.750 1.0141 0.02114 0.01199 -0.0663 0.4244 1.0000
6.000 1.0393 0.02167 0.01257 -0.0657 0.4133 1.0000
6.250 1.0648 0.02217 0.01306 -0.0652 0.4023 1.0000
6.500 1.0885 0.02275 0.01378 -0.0646 0.3901 1.0000
6.750 1.1121 0.02305 0.01415 -0.0638 0.3763 1.0000
7.000 1.1354 0.02301 0.01413 -0.0628 0.3604 1.0000
7.250 1.1582 0.02304 0.01419 -0.0618 0.3452 1.0000
7.500 1.1805 0.02307 0.01428 -0.0608 0.3301 1.0000
7.750 1.2016 0.02313 0.01447 -0.0597 0.3143 1.0000
8.000 1.2213 0.02310 0.01462 -0.0584 0.2945 1.0000
8.250 1.2415 0.02322 0.01490 -0.0571 0.2763 1.0000
8.500 1.2589 0.02341 0.01513 -0.0556 0.2474 1.0000
8.750 1.2762 0.02401 0.01573 -0.0542 0.2151 1.0000
9.000 1.2907 0.02510 0.01669 -0.0526 0.1672 1.0000
9.250 1.2891 0.02795 0.01874 -0.0499 0.0716 1.0000
9.500 1.2916 0.03036 0.02101 -0.0473 0.0542 1.0000
9.750 1.2968 0.03230 0.02306 -0.0450 0.0485 1.0000
10.000 1.2967 0.03439 0.02531 -0.0423 0.0452 1.0000
10.500 1.2884 0.03924 0.03051 -0.0378 0.0424 1.0000
10.750 1.2839 0.04205 0.03352 -0.0366 0.0417 1.0000
11.000 1.2783 0.04526 0.03692 -0.0360 0.0412 1.0000
11.250 1.2726 0.04878 0.04061 -0.0360 0.0407 1.0000
11.500 1.2676 0.05237 0.04434 -0.0361 0.0403 1.0000
11.750 1.2644 0.05576 0.04785 -0.0360 0.0398 1.0000
12.000 1.2644 0.05868 0.05086 -0.0353 0.0394 1.0000
12.250 1.2691 0.06091 0.05315 -0.0339 0.0390 1.0000
12.500 1.2779 0.06270 0.05498 -0.0320 0.0382 1.0000
12.750 1.2880 0.06453 0.05683 -0.0301 0.0370 1.0000
13.000 1.3030 0.06608 0.05841 -0.0277 0.0357 1.0000
13.250 1.3222 0.06787 0.06029 -0.0251 0.0353 1.0000
13.500 1.3320 0.07070 0.06334 -0.0239 0.0355 1.0000
13.750 1.3344 0.07425 0.06717 -0.0235 0.0359 1.0000
14.000 1.3309 0.07840 0.07163 -0.0237 0.0363 1.0000
14.250 1.3235 0.08309 0.07661 -0.0245 0.0369 1.0000
14.500 1.3081 0.08865 0.08248 -0.0264 0.0374 1.0000
14.750 1.2898 0.09480 0.08891 -0.0290 0.0378 1.0000
15.000 1.2698 0.10145 0.09583 -0.0324 0.0383 1.0000
15.250 1.2484 0.10871 0.10333 -0.0364 0.0387 1.0000
15.500 1.2259 0.11668 0.11152 -0.0413 0.0391 1.0000
15.750 1.2021 0.12550 0.12054 -0.0470 0.0395 1.0000
16.000 1.1767 0.13533 0.13057 -0.0537 0.0399 1.0000
16.250 1.1490 0.14669 0.14207 -0.0616 0.0404 1.0000
16.500 1.1197 0.15971 0.15517 -0.0705 0.0412 1.0000
16.750 1.0985 0.17157 0.16705 -0.0778 0.0423 1.0000
17.000 1.0864 0.18199 0.17749 -0.0836 0.0439 1.0000
17.250 0.7919 0.17347 0.16915 -0.0599 0.0521 1.0000
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