USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 200,000 Max Cl/Cd: 72.97 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa22-il-200000.txt Download as CSV file: xf-usa22-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3856 0.11099 0.10758 -0.0113 1.0000 0.0385
-9.000 -0.3837 0.10824 0.10487 -0.0151 1.0000 0.0389
-8.750 -0.3808 0.10510 0.10178 -0.0183 1.0000 0.0391
-8.500 -0.3768 0.10179 0.09851 -0.0209 1.0000 0.0392
-8.250 -0.3734 0.09836 0.09513 -0.0234 1.0000 0.0392
-8.000 -0.3682 0.09208 0.08888 -0.0205 1.0000 0.0404
-7.750 -0.3591 0.08907 0.08589 -0.0196 1.0000 0.0414
-7.500 -0.3528 0.08621 0.08307 -0.0199 1.0000 0.0424
-7.250 -0.3481 0.08328 0.08018 -0.0211 1.0000 0.0435
-7.000 -0.3419 0.08005 0.07699 -0.0233 1.0000 0.0449
-6.750 -0.3358 0.07670 0.07367 -0.0263 1.0000 0.0470
-6.500 -0.3261 0.07259 0.06958 -0.0351 1.0000 0.0503
-6.250 -0.3144 0.06856 0.06549 -0.0416 1.0000 0.0510
-6.000 -0.3181 0.06275 0.05970 -0.0420 1.0000 0.0519
-5.750 -0.3188 0.06086 0.05788 -0.0381 1.0000 0.0530
-5.500 -0.3135 0.05880 0.05585 -0.0366 0.9995 0.0541
-5.250 -0.2744 0.05444 0.05140 -0.0432 0.9940 0.0575
-5.000 -0.1641 0.02847 0.02534 -0.0629 0.9668 0.0683
-4.750 -0.1847 0.04141 0.03787 -0.0616 0.9789 0.0687
-4.500 -0.1351 0.02837 0.02360 -0.0703 0.9720 0.0490
-4.250 -0.0959 0.02323 0.01775 -0.0736 0.9639 0.0470
-4.000 -0.0532 0.02014 0.01400 -0.0764 0.9571 0.0486
-3.750 -0.0164 0.01794 0.01132 -0.0780 0.9457 0.0536
-3.500 0.0200 0.01647 0.00968 -0.0794 0.9330 0.0582
-3.250 0.0554 0.01588 0.00879 -0.0801 0.9177 0.0646
-3.000 0.0863 0.01422 0.00708 -0.0804 0.9009 0.0711
-2.750 0.1151 0.01368 0.00638 -0.0799 0.8803 0.0779
-2.500 0.1411 0.01282 0.00549 -0.0791 0.8576 0.0844
-2.250 0.1672 0.01243 0.00497 -0.0781 0.8334 0.0921
-2.000 0.1927 0.01192 0.00443 -0.0771 0.8092 0.1013
-1.750 0.2184 0.01160 0.00406 -0.0762 0.7839 0.1188
-1.500 0.2436 0.01108 0.00381 -0.0754 0.7591 0.1964
-1.250 0.2697 0.01116 0.00387 -0.0746 0.7351 0.2773
-1.000 0.2959 0.01129 0.00384 -0.0740 0.7116 0.3101
-0.750 0.3221 0.01138 0.00380 -0.0734 0.6890 0.3364
-0.500 0.3481 0.01137 0.00372 -0.0728 0.6677 0.3592
-0.250 0.3740 0.01134 0.00362 -0.0723 0.6480 0.3846
0.000 0.3999 0.01122 0.00352 -0.0718 0.6288 0.4139
0.250 0.4245 0.01088 0.00343 -0.0712 0.6117 0.4791
0.500 0.4615 0.00985 0.00339 -0.0724 0.5946 1.0000
0.750 0.4882 0.01004 0.00337 -0.0719 0.5795 1.0000
1.000 0.5149 0.01023 0.00338 -0.0716 0.5652 1.0000
1.250 0.5415 0.01041 0.00341 -0.0712 0.5515 1.0000
1.500 0.5682 0.01060 0.00346 -0.0708 0.5386 1.0000
1.750 0.5948 0.01080 0.00353 -0.0705 0.5268 1.0000
2.000 0.6214 0.01102 0.00361 -0.0701 0.5160 1.0000
2.250 0.6481 0.01122 0.00372 -0.0698 0.5055 1.0000
2.500 0.6749 0.01142 0.00386 -0.0695 0.4953 1.0000
2.750 0.7016 0.01165 0.00400 -0.0693 0.4862 1.0000
3.000 0.7283 0.01187 0.00416 -0.0690 0.4770 1.0000
3.250 0.7549 0.01209 0.00437 -0.0687 0.4677 1.0000
3.500 0.7814 0.01236 0.00454 -0.0684 0.4590 1.0000
3.750 0.8079 0.01256 0.00474 -0.0681 0.4492 1.0000
4.000 0.8343 0.01281 0.00497 -0.0678 0.4401 1.0000
4.250 0.8607 0.01309 0.00520 -0.0675 0.4316 1.0000
4.500 0.8870 0.01330 0.00547 -0.0672 0.4218 1.0000
4.750 0.9131 0.01358 0.00573 -0.0669 0.4127 1.0000
5.000 0.9392 0.01384 0.00599 -0.0666 0.4031 1.0000
5.250 0.9650 0.01408 0.00629 -0.0662 0.3923 1.0000
5.500 0.9907 0.01435 0.00658 -0.0658 0.3817 1.0000
5.750 1.0160 0.01459 0.00682 -0.0654 0.3700 1.0000
6.000 1.0406 0.01471 0.00699 -0.0648 0.3538 1.0000
6.250 1.0653 0.01487 0.00719 -0.0643 0.3378 1.0000
6.500 1.0900 0.01510 0.00745 -0.0638 0.3234 1.0000
6.750 1.1140 0.01530 0.00764 -0.0632 0.3039 1.0000
7.000 1.1378 0.01560 0.00796 -0.0626 0.2866 1.0000
7.250 1.1617 0.01592 0.00832 -0.0620 0.2665 1.0000
7.500 1.1833 0.01643 0.00869 -0.0613 0.2338 1.0000
7.750 1.2042 0.01712 0.00923 -0.0604 0.1973 1.0000
8.000 1.2222 0.01822 0.01001 -0.0592 0.1434 1.0000
8.250 1.2243 0.02119 0.01214 -0.0565 0.0371 1.0000
8.500 1.2419 0.02229 0.01336 -0.0552 0.0311 1.0000
8.750 1.2584 0.02344 0.01469 -0.0537 0.0285 1.0000
9.000 1.2722 0.02479 0.01623 -0.0519 0.0269 1.0000
9.250 1.2814 0.02640 0.01807 -0.0498 0.0258 1.0000
9.500 1.2877 0.02807 0.01990 -0.0474 0.0253 1.0000
9.750 1.2915 0.02963 0.02160 -0.0447 0.0250 1.0000
10.000 1.2934 0.03134 0.02347 -0.0421 0.0246 1.0000
10.250 1.2946 0.03327 0.02551 -0.0400 0.0240 1.0000
10.500 1.2950 0.03545 0.02780 -0.0384 0.0233 1.0000
10.750 1.2952 0.03785 0.03030 -0.0371 0.0227 1.0000
11.000 1.2959 0.04038 0.03292 -0.0360 0.0225 1.0000
11.250 1.2981 0.04289 0.03552 -0.0349 0.0223 1.0000
11.500 1.3021 0.04528 0.03800 -0.0337 0.0221 1.0000
11.750 1.3084 0.04753 0.04032 -0.0324 0.0221 1.0000
12.000 1.3171 0.04964 0.04252 -0.0308 0.0221 1.0000
12.250 1.3272 0.05180 0.04480 -0.0292 0.0224 1.0000
12.500 1.3387 0.05409 0.04725 -0.0275 0.0228 1.0000
12.750 1.3503 0.05673 0.05007 -0.0258 0.0235 1.0000
13.000 1.3576 0.05984 0.05338 -0.0245 0.0240 1.0000
13.250 1.3595 0.06334 0.05708 -0.0238 0.0242 1.0000
13.500 1.3565 0.06711 0.06107 -0.0235 0.0243 1.0000
13.750 1.3510 0.07128 0.06546 -0.0235 0.0244 1.0000
14.500 1.3578 0.08379 0.07894 -0.0221 0.0356 1.0000
14.750 1.3285 0.09017 0.08565 -0.0255 0.0363 1.0000
15.000 1.3017 0.09711 0.09284 -0.0295 0.0367 1.0000
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Polar data table (+)
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