USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: USA 22 AIRFOIL (usa22-il) Reynolds number: 500,000 Max Cl/Cd: 99.94 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa22-il-500000.txt Download as CSV file: xf-usa22-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2809 0.09367 0.09164 -0.0193 1.0000 0.0215
-9.000 -0.2836 0.08961 0.08760 -0.0220 1.0000 0.0221
-8.750 -0.2824 0.08573 0.08375 -0.0235 1.0000 0.0222
-8.500 -0.2809 0.08182 0.07985 -0.0249 1.0000 0.0223
-8.250 -0.2800 0.07783 0.07589 -0.0262 1.0000 0.0223
-8.000 -0.2801 0.07382 0.07190 -0.0275 1.0000 0.0224
-7.750 -0.2815 0.06984 0.06795 -0.0288 1.0000 0.0224
-7.500 -0.2953 0.06409 0.06226 -0.0289 1.0000 0.0230
-7.250 -0.3013 0.06147 0.05970 -0.0279 1.0000 0.0232
-7.000 -0.3000 0.05840 0.05666 -0.0287 0.9986 0.0235
-6.750 -0.2803 0.05370 0.05195 -0.0344 0.9930 0.0240
-6.500 -0.2594 0.04843 0.04665 -0.0411 0.9855 0.0248
-6.250 -0.2976 0.03097 0.02818 -0.0672 0.9865 0.0166
-6.000 -0.2678 0.02147 0.01759 -0.0729 0.9796 0.0167
-5.750 -0.2366 0.01745 0.01298 -0.0752 0.9728 0.0179
-5.500 -0.2033 0.01696 0.01241 -0.0765 0.9645 0.0192
-5.250 -0.1697 0.01541 0.01057 -0.0777 0.9557 0.0204
-5.000 -0.1385 0.01418 0.00907 -0.0782 0.9430 0.0217
-4.750 -0.1094 0.01267 0.00727 -0.0783 0.9280 0.0233
-4.500 -0.0816 0.01230 0.00684 -0.0781 0.9105 0.0254
-4.250 -0.0546 0.01211 0.00654 -0.0775 0.8916 0.0279
-4.000 -0.0280 0.01182 0.00608 -0.0768 0.8710 0.0299
-3.750 -0.0027 0.01096 0.00510 -0.0762 0.8479 0.0337
-3.500 0.0238 0.01089 0.00489 -0.0755 0.8218 0.0375
-3.250 0.0495 0.01039 0.00421 -0.0748 0.7944 0.0416
-3.000 0.0756 0.01017 0.00386 -0.0742 0.7670 0.0456
-2.750 0.1023 0.01011 0.00364 -0.0737 0.7415 0.0496
-2.500 0.1287 0.00982 0.00319 -0.0733 0.7192 0.0532
-2.250 0.1553 0.00956 0.00284 -0.0729 0.6985 0.0584
-2.000 0.1823 0.00945 0.00259 -0.0725 0.6787 0.0626
-1.750 0.2094 0.00928 0.00231 -0.0722 0.6585 0.0676
-1.500 0.2364 0.00912 0.00208 -0.0718 0.6385 0.0754
-1.000 0.2897 0.00860 0.00184 -0.0713 0.5976 0.2016
-0.750 0.3167 0.00862 0.00186 -0.0711 0.5774 0.2503
-0.500 0.3440 0.00868 0.00187 -0.0708 0.5576 0.2753
-0.250 0.3713 0.00877 0.00187 -0.0706 0.5390 0.2941
0.000 0.3986 0.00884 0.00189 -0.0704 0.5218 0.3122
0.250 0.4258 0.00891 0.00189 -0.0702 0.5060 0.3272
0.500 0.4532 0.00897 0.00190 -0.0701 0.4917 0.3405
0.750 0.4805 0.00901 0.00192 -0.0699 0.4794 0.3582
1.000 0.5079 0.00903 0.00194 -0.0698 0.4693 0.3776
1.500 0.5670 0.00765 0.00205 -0.0705 0.4519 1.0000
1.750 0.5942 0.00780 0.00210 -0.0703 0.4440 1.0000
2.000 0.6215 0.00791 0.00215 -0.0701 0.4359 1.0000
2.250 0.6486 0.00806 0.00223 -0.0699 0.4288 1.0000
2.500 0.6760 0.00818 0.00231 -0.0697 0.4216 1.0000
2.750 0.7030 0.00834 0.00240 -0.0695 0.4151 1.0000
3.000 0.7304 0.00845 0.00249 -0.0694 0.4081 1.0000
3.250 0.7573 0.00862 0.00261 -0.0691 0.4015 1.0000
3.500 0.7848 0.00872 0.00272 -0.0690 0.3945 1.0000
4.000 0.8390 0.00900 0.00297 -0.0687 0.3796 1.0000
4.250 0.8657 0.00918 0.00310 -0.0685 0.3719 1.0000
4.500 0.8930 0.00929 0.00324 -0.0683 0.3636 1.0000
4.750 0.9198 0.00945 0.00338 -0.0681 0.3544 1.0000
5.000 0.9465 0.00961 0.00352 -0.0679 0.3407 1.0000
5.250 0.9730 0.00979 0.00366 -0.0677 0.3257 1.0000
5.500 0.9994 0.01000 0.00383 -0.0675 0.3098 1.0000
5.750 1.0253 0.01026 0.00403 -0.0672 0.2917 1.0000
6.000 1.0507 0.01058 0.00428 -0.0669 0.2730 1.0000
6.250 1.0754 0.01098 0.00457 -0.0665 0.2467 1.0000
6.500 1.0993 0.01150 0.00492 -0.0661 0.2180 1.0000
6.750 1.1225 0.01210 0.00535 -0.0655 0.1838 1.0000
7.000 1.1380 0.01373 0.00633 -0.0642 0.0871 1.0000
7.250 1.1531 0.01540 0.00767 -0.0626 0.0225 1.0000
7.500 1.1757 0.01608 0.00845 -0.0618 0.0186 1.0000
7.750 1.1986 0.01666 0.00914 -0.0611 0.0176 1.0000
8.000 1.2205 0.01735 0.00994 -0.0602 0.0166 1.0000
8.250 1.2411 0.01817 0.01087 -0.0592 0.0157 1.0000
8.500 1.2599 0.01912 0.01191 -0.0581 0.0147 1.0000
8.750 1.2742 0.02049 0.01340 -0.0564 0.0136 1.0000
9.000 1.2804 0.02249 0.01557 -0.0538 0.0129 1.0000
9.250 1.2968 0.02339 0.01656 -0.0524 0.0127 1.0000
9.500 1.3099 0.02452 0.01778 -0.0506 0.0124 1.0000
9.750 1.3193 0.02579 0.01916 -0.0484 0.0121 1.0000
10.000 1.3244 0.02718 0.02064 -0.0457 0.0119 1.0000
10.250 1.3278 0.02877 0.02233 -0.0432 0.0117 1.0000
10.500 1.3301 0.03059 0.02425 -0.0411 0.0115 1.0000
10.750 1.3323 0.03262 0.02636 -0.0392 0.0113 1.0000
11.000 1.3352 0.03474 0.02857 -0.0378 0.0112 1.0000
11.250 1.3389 0.03690 0.03082 -0.0366 0.0109 1.0000
11.500 1.3432 0.03908 0.03308 -0.0357 0.0106 1.0000
11.750 1.3474 0.04134 0.03539 -0.0350 0.0103 1.0000
12.000 1.3511 0.04371 0.03783 -0.0344 0.0100 1.0000
12.250 1.3553 0.04610 0.04029 -0.0335 0.0098 1.0000
12.500 1.3605 0.04844 0.04271 -0.0324 0.0097 1.0000
12.750 1.3665 0.05080 0.04514 -0.0312 0.0097 1.0000
13.000 1.3727 0.05321 0.04766 -0.0301 0.0097 1.0000
13.250 1.3732 0.05673 0.05149 -0.0288 0.0103 1.0000
13.500 1.1908 0.05920 0.05442 -0.0214 0.0102 1.0000
13.750 1.1884 0.06306 0.05840 -0.0206 0.0103 1.0000
14.000 1.1827 0.06733 0.06285 -0.0199 0.0105 1.0000
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Polar data table (+)
Polar graphs
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