USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: USA 5 AIRFOIL (usa5-il) Reynolds number: 500,000 Max Cl/Cd: 94.96 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-usa5-il-500000.txt Download as CSV file: xf-usa5-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: USA 5 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3529 0.09098 0.08897 -0.0212 1.0000 0.0127
-7.250 -0.3613 0.08925 0.08728 -0.0191 1.0000 0.0127
-7.000 -0.3666 0.08702 0.08508 -0.0182 1.0000 0.0127
-6.750 -0.3550 0.08331 0.08137 -0.0217 0.9985 0.0127
-6.500 -0.3415 0.07684 0.07491 -0.0257 0.9956 0.0137
-6.250 -0.3152 0.07273 0.07076 -0.0315 0.9922 0.0142
-6.000 -0.2869 0.06865 0.06664 -0.0377 0.9880 0.0160
-5.750 -0.2520 0.06367 0.06161 -0.0461 0.9844 0.0175
-5.500 -0.2062 0.05923 0.05706 -0.0558 0.9807 0.0192
-5.250 -0.1718 0.05479 0.05252 -0.0617 0.9752 0.0194
-5.000 -0.1454 0.04673 0.04434 -0.0686 0.9714 0.0205
-4.750 -0.1125 0.04340 0.04093 -0.0726 0.9693 0.0213
-4.500 -0.0851 0.04048 0.03792 -0.0749 0.9634 0.0223
-4.250 -0.0509 0.03689 0.03417 -0.0784 0.9592 0.0237
-4.000 -0.0135 0.03274 0.02981 -0.0819 0.9554 0.0262
-3.750 0.0009 0.01915 0.01523 -0.0795 0.9444 0.0163
-3.500 0.0228 0.01507 0.01043 -0.0779 0.9366 0.0172
-3.250 0.0497 0.01324 0.00819 -0.0772 0.9293 0.0180
-3.000 0.0775 0.01310 0.00793 -0.0768 0.9207 0.0194
-2.750 0.1033 0.01069 0.00509 -0.0761 0.9134 0.0218
-2.500 0.1284 0.01011 0.00443 -0.0752 0.9025 0.0246
-2.250 0.1549 0.00961 0.00381 -0.0745 0.8910 0.0277
-2.000 0.1805 0.00898 0.00308 -0.0737 0.8777 0.0343
-1.750 0.2068 0.00870 0.00268 -0.0730 0.8617 0.0408
-1.250 0.2579 0.00824 0.00217 -0.0714 0.8244 0.0718
-1.000 0.2845 0.00836 0.00220 -0.0709 0.8034 0.0894
-0.750 0.3100 0.00837 0.00212 -0.0702 0.7819 0.0996
-0.500 0.3351 0.00840 0.00202 -0.0694 0.7586 0.1078
-0.250 0.3593 0.00840 0.00192 -0.0685 0.7329 0.1146
0.000 0.3835 0.00848 0.00183 -0.0675 0.7056 0.1190
0.250 0.4066 0.00843 0.00169 -0.0664 0.6776 0.1252
0.500 0.4304 0.00852 0.00165 -0.0655 0.6505 0.1314
0.750 0.4538 0.00855 0.00159 -0.0644 0.6235 0.1387
1.000 0.4771 0.00865 0.00157 -0.0634 0.5937 0.1449
1.250 0.5002 0.00874 0.00155 -0.0624 0.5579 0.1518
1.500 0.5229 0.00888 0.00156 -0.0613 0.5186 0.1623
1.750 0.5451 0.00901 0.00161 -0.0601 0.4811 0.1899
2.000 0.6663 0.00778 0.00195 -0.0825 0.4321 1.0000
2.250 0.6906 0.00798 0.00204 -0.0817 0.4175 1.0000
2.500 0.7151 0.00816 0.00214 -0.0810 0.4059 1.0000
2.750 0.7397 0.00835 0.00225 -0.0804 0.3962 1.0000
3.000 0.7643 0.00852 0.00237 -0.0797 0.3870 1.0000
3.250 0.7892 0.00867 0.00252 -0.0791 0.3794 1.0000
3.500 0.8135 0.00887 0.00266 -0.0784 0.3704 1.0000
3.750 0.8379 0.00904 0.00279 -0.0777 0.3572 1.0000
4.000 0.8623 0.00922 0.00292 -0.0771 0.3434 1.0000
4.250 0.8865 0.00940 0.00308 -0.0764 0.3291 1.0000
4.500 0.9107 0.00959 0.00322 -0.0757 0.3094 1.0000
4.750 0.9337 0.00988 0.00337 -0.0749 0.2751 1.0000
5.000 0.9504 0.01080 0.00373 -0.0731 0.1776 1.0000
5.250 0.9698 0.01150 0.00422 -0.0717 0.1403 1.0000
5.500 0.9814 0.01305 0.00514 -0.0690 0.0232 1.0000
5.750 1.0034 0.01351 0.00568 -0.0679 0.0183 1.0000
6.000 1.0252 0.01397 0.00626 -0.0666 0.0162 1.0000
6.250 1.0458 0.01458 0.00701 -0.0652 0.0148 1.0000
6.500 1.0644 0.01537 0.00794 -0.0634 0.0142 1.0000
6.750 1.0813 0.01627 0.00897 -0.0613 0.0135 1.0000
7.000 1.1006 0.01689 0.00966 -0.0598 0.0127 1.0000
7.250 1.1168 0.01777 0.01066 -0.0577 0.0122 1.0000
7.500 1.1311 0.01878 0.01178 -0.0553 0.0120 1.0000
7.750 1.1439 0.01992 0.01301 -0.0526 0.0116 1.0000
8.000 1.1550 0.02123 0.01442 -0.0497 0.0113 1.0000
8.250 1.1657 0.02273 0.01601 -0.0467 0.0116 1.0000
8.500 1.1775 0.02464 0.01802 -0.0439 0.0124 1.0000
14.000 0.9359 0.09906 0.09630 -0.0096 0.0136 1.0000
14.250 0.9118 0.10526 0.10260 -0.0128 0.0136 1.0000
14.500 0.8858 0.11073 0.10818 -0.0162 0.0136 1.0000
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