USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 200,000 Max Cl/Cd: 73.53 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa22-il-200000-n5.txt Download as CSV file: xf-usa22-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2634 0.09215 0.08901 -0.0235 1.0000 0.0152
-9.000 -0.3647 0.10266 0.09930 -0.0146 1.0000 0.0153
-8.750 -0.3594 0.09925 0.09592 -0.0156 1.0000 0.0149
-8.500 -0.3549 0.09577 0.09247 -0.0170 1.0000 0.0146
-8.250 -0.3511 0.09224 0.08898 -0.0185 1.0000 0.0144
-8.000 -0.3481 0.08869 0.08547 -0.0202 1.0000 0.0142
-7.750 -0.3464 0.08518 0.08201 -0.0218 1.0000 0.0140
-7.500 -0.3466 0.08175 0.07865 -0.0232 1.0000 0.0138
-7.250 -0.3447 0.07795 0.07490 -0.0257 1.0000 0.0137
-7.000 -0.3426 0.07399 0.07101 -0.0285 1.0000 0.0136
-6.750 -0.3408 0.07002 0.06709 -0.0310 1.0000 0.0137
-6.500 -0.3143 0.06295 0.05994 -0.0413 0.9893 0.0143
-6.000 -0.2491 0.04321 0.03984 -0.0654 0.9622 0.0150
-5.750 -0.2175 0.02580 0.02109 -0.0773 0.9488 0.0152
-5.500 -0.1856 0.02136 0.01579 -0.0796 0.9379 0.0159
-5.250 -0.1546 0.01907 0.01305 -0.0809 0.9254 0.0175
-5.000 -0.1225 0.01842 0.01223 -0.0819 0.9117 0.0198
-4.750 -0.0919 0.01690 0.01031 -0.0823 0.8967 0.0222
-4.500 -0.0626 0.01564 0.00873 -0.0823 0.8804 0.0247
-4.250 -0.0344 0.01535 0.00836 -0.0822 0.8613 0.0287
-4.000 -0.0066 0.01474 0.00747 -0.0818 0.8406 0.0332
-3.750 0.0202 0.01432 0.00696 -0.0814 0.8175 0.0383
-3.500 0.0473 0.01404 0.00644 -0.0808 0.7937 0.0434
-3.250 0.0734 0.01352 0.00579 -0.0803 0.7704 0.0489
-3.000 0.0999 0.01327 0.00535 -0.0798 0.7481 0.0544
-2.750 0.1265 0.01306 0.00491 -0.0792 0.7267 0.0579
-2.500 0.1526 0.01262 0.00435 -0.0787 0.7073 0.0618
-2.250 0.1793 0.01240 0.00397 -0.0783 0.6890 0.0660
-2.000 0.2061 0.01225 0.00367 -0.0779 0.6706 0.0710
-1.750 0.2329 0.01205 0.00339 -0.0775 0.6524 0.0802
-1.500 0.2596 0.01186 0.00317 -0.0772 0.6348 0.0965
-1.250 0.2861 0.01163 0.00303 -0.0769 0.6177 0.1532
-1.000 0.3127 0.01152 0.00298 -0.0766 0.6005 0.2024
-0.750 0.3395 0.01152 0.00297 -0.0763 0.5836 0.2448
-0.500 0.3664 0.01157 0.00297 -0.0760 0.5676 0.2761
-0.250 0.3933 0.01162 0.00295 -0.0757 0.5523 0.2978
0.000 0.4202 0.01168 0.00292 -0.0754 0.5379 0.3171
0.250 0.4469 0.01169 0.00292 -0.0752 0.5246 0.3381
0.500 0.4736 0.01172 0.00291 -0.0749 0.5121 0.3608
0.750 0.5002 0.01172 0.00289 -0.0746 0.5002 0.3802
1.250 0.5486 0.01097 0.00295 -0.0732 0.4778 0.6956
1.750 0.6120 0.01082 0.00302 -0.0743 0.4553 1.0000
2.000 0.6389 0.01099 0.00309 -0.0741 0.4452 1.0000
2.250 0.6656 0.01117 0.00319 -0.0738 0.4365 1.0000
2.500 0.6923 0.01136 0.00330 -0.0736 0.4283 1.0000
2.750 0.7191 0.01153 0.00343 -0.0733 0.4208 1.0000
3.000 0.7457 0.01173 0.00357 -0.0731 0.4132 1.0000
3.250 0.7724 0.01191 0.00373 -0.0729 0.4055 1.0000
3.500 0.7988 0.01212 0.00389 -0.0726 0.3978 1.0000
3.750 0.8255 0.01230 0.00408 -0.0724 0.3901 1.0000
4.000 0.8519 0.01252 0.00427 -0.0721 0.3828 1.0000
4.250 0.8785 0.01270 0.00448 -0.0719 0.3751 1.0000
4.500 0.9046 0.01293 0.00471 -0.0716 0.3675 1.0000
4.750 0.9311 0.01312 0.00495 -0.0714 0.3591 1.0000
5.000 0.9571 0.01336 0.00519 -0.0711 0.3506 1.0000
5.250 0.9831 0.01358 0.00547 -0.0708 0.3408 1.0000
5.500 1.0090 0.01382 0.00575 -0.0705 0.3310 1.0000
5.750 1.0343 0.01408 0.00602 -0.0701 0.3174 1.0000
6.000 1.0588 0.01440 0.00628 -0.0696 0.2968 1.0000
6.250 1.0831 0.01476 0.00663 -0.0692 0.2776 1.0000
6.500 1.1064 0.01521 0.00700 -0.0686 0.2537 1.0000
6.750 1.1283 0.01584 0.00747 -0.0679 0.2226 1.0000
7.000 1.1496 0.01655 0.00803 -0.0672 0.1905 1.0000
7.250 1.1692 0.01748 0.00873 -0.0663 0.1491 1.0000
7.500 1.1720 0.02041 0.01081 -0.0637 0.0241 1.0000
7.750 1.1915 0.02133 0.01184 -0.0625 0.0185 1.0000
8.000 1.2116 0.02211 0.01279 -0.0615 0.0168 1.0000
8.250 1.2307 0.02296 0.01381 -0.0604 0.0150 1.0000
8.500 1.2481 0.02395 0.01498 -0.0591 0.0136 1.0000
8.750 1.2632 0.02510 0.01632 -0.0575 0.0128 1.0000
9.000 1.2748 0.02646 0.01788 -0.0557 0.0121 1.0000
9.250 1.2810 0.02813 0.01975 -0.0534 0.0115 1.0000
9.500 1.2775 0.03017 0.02198 -0.0500 0.0110 1.0000
9.750 1.2750 0.03224 0.02421 -0.0474 0.0106 1.0000
10.000 1.2805 0.03386 0.02596 -0.0458 0.0101 1.0000
10.250 1.2822 0.03596 0.02819 -0.0445 0.0097 1.0000
10.500 1.2809 0.03855 0.03090 -0.0435 0.0095 1.0000
10.750 1.2793 0.04142 0.03390 -0.0430 0.0093 1.0000
11.000 1.2776 0.04449 0.03708 -0.0427 0.0091 1.0000
11.250 1.2761 0.04763 0.04033 -0.0426 0.0089 1.0000
11.500 1.2750 0.05079 0.04358 -0.0425 0.0087 1.0000
11.750 1.2744 0.05390 0.04679 -0.0423 0.0086 1.0000
12.000 1.2743 0.05693 0.04991 -0.0421 0.0084 1.0000
12.250 1.2751 0.05987 0.05293 -0.0417 0.0083 1.0000
12.500 1.2766 0.06270 0.05584 -0.0413 0.0081 1.0000
12.750 1.2784 0.06553 0.05875 -0.0408 0.0080 1.0000
13.000 1.2797 0.06849 0.06179 -0.0406 0.0078 1.0000
13.250 1.2804 0.07151 0.06489 -0.0405 0.0076 1.0000
13.500 1.2814 0.07443 0.06785 -0.0400 0.0073 1.0000
13.750 1.2833 0.07739 0.07088 -0.0390 0.0070 1.0000
14.000 1.2781 0.08160 0.07535 -0.0405 0.0069 1.0000
14.250 1.2737 0.08582 0.07977 -0.0417 0.0067 1.0000
14.500 1.2697 0.09004 0.08418 -0.0428 0.0066 1.0000
14.750 1.2651 0.09447 0.08878 -0.0440 0.0065 1.0000
15.000 1.2592 0.09918 0.09367 -0.0455 0.0065 1.0000
15.250 1.2522 0.10422 0.09889 -0.0474 0.0064 1.0000
15.500 1.2446 0.10948 0.10433 -0.0495 0.0064 1.0000
15.750 1.2361 0.11509 0.11012 -0.0521 0.0064 1.0000
16.000 1.2271 0.12095 0.11616 -0.0549 0.0064 1.0000
16.250 1.2175 0.12709 0.12247 -0.0581 0.0064 1.0000
16.500 1.2079 0.13344 0.12898 -0.0616 0.0064 1.0000
16.750 1.1980 0.14013 0.13584 -0.0655 0.0064 1.0000
17.000 1.1881 0.14707 0.14293 -0.0697 0.0064 1.0000
17.250 1.1781 0.15429 0.15030 -0.0742 0.0064 1.0000
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