NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il) Reynolds number: 500,000 Max Cl/Cd: 67.04 at α=0.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20706-il-500000.txt Download as CSV file: xf-sc20706-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0706 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4914 0.08303 0.08082 -0.0184 1.0000 0.0184
-9.750 -0.4957 0.07867 0.07648 -0.0192 1.0000 0.0187
-9.500 -0.5016 0.07421 0.07205 -0.0200 1.0000 0.0190
-9.250 -0.5092 0.06967 0.06753 -0.0209 1.0000 0.0193
-9.000 -0.5196 0.06515 0.06305 -0.0216 1.0000 0.0195
-8.750 -0.5348 0.06089 0.05882 -0.0218 1.0000 0.0194
-8.500 -0.5539 0.05564 0.05362 -0.0246 1.0000 0.0190
-8.250 -0.5661 0.04854 0.04643 -0.0320 1.0000 0.0188
-8.000 -0.5679 0.04259 0.04035 -0.0358 1.0000 0.0192
-7.750 -0.5622 0.03692 0.03451 -0.0390 1.0000 0.0200
-7.500 -0.5420 0.03355 0.03083 -0.0407 1.0000 0.0230
-7.250 -0.5404 0.04154 0.03820 -0.0471 1.0000 0.0234
-6.250 -0.4015 0.01640 0.01077 -0.0631 1.0000 0.0175
-6.000 -0.3698 0.01486 0.00901 -0.0642 1.0000 0.0182
-5.750 -0.3388 0.01375 0.00779 -0.0650 1.0000 0.0193
-5.500 -0.3074 0.01283 0.00677 -0.0660 1.0000 0.0205
-5.250 -0.2773 0.01230 0.00617 -0.0666 1.0000 0.0224
-5.000 -0.2426 0.01142 0.00518 -0.0684 1.0000 0.0243
-4.750 -0.2046 0.01047 0.00417 -0.0710 1.0000 0.0299
-4.500 -0.1717 0.01006 0.00369 -0.0723 1.0000 0.0360
-4.250 -0.1383 0.00963 0.00330 -0.0738 1.0000 0.0513
-4.000 -0.0973 0.00876 0.00292 -0.0774 1.0000 0.1729
-3.750 -0.0458 0.00742 0.00266 -0.0840 1.0000 0.4801
-3.500 -0.0123 0.00716 0.00271 -0.0855 1.0000 0.5844
-3.250 0.0171 0.00711 0.00275 -0.0859 1.0000 0.6238
-3.000 0.0456 0.00712 0.00279 -0.0860 1.0000 0.6519
-2.750 0.0736 0.00715 0.00288 -0.0860 1.0000 0.6799
-2.500 0.1013 0.00720 0.00296 -0.0859 1.0000 0.6998
-2.250 0.1285 0.00726 0.00307 -0.0857 1.0000 0.7172
-2.000 0.1558 0.00732 0.00317 -0.0855 1.0000 0.7305
-1.750 0.1829 0.00740 0.00329 -0.0853 1.0000 0.7436
-1.500 0.2098 0.00749 0.00342 -0.0851 1.0000 0.7553
-1.250 0.2367 0.00756 0.00355 -0.0848 1.0000 0.7642
-1.000 0.2640 0.00764 0.00369 -0.0848 1.0000 0.7725
-0.750 0.2910 0.00772 0.00383 -0.0846 1.0000 0.7797
-0.500 0.3181 0.00781 0.00399 -0.0845 1.0000 0.7869
-0.250 0.3902 0.00707 0.00336 -0.0939 0.9811 0.7922
0.000 0.4355 0.00666 0.00304 -0.0973 0.9645 0.7986
0.250 0.4887 0.00729 0.00246 -0.1016 0.6233 0.8022
0.500 0.5047 0.00914 0.00289 -0.0997 0.2786 0.8083
0.750 0.5269 0.01046 0.00330 -0.0990 0.0670 0.8142
1.000 0.5527 0.01098 0.00377 -0.0984 0.0419 0.8198
1.250 0.5796 0.01146 0.00426 -0.0981 0.0316 0.8263
1.500 0.6053 0.01189 0.00477 -0.0974 0.0274 0.8316
1.750 0.6317 0.01241 0.00534 -0.0970 0.0239 0.8382
2.000 0.6549 0.01368 0.00670 -0.0957 0.0211 0.8436
2.250 0.6792 0.01508 0.00825 -0.0946 0.0202 0.8497
2.500 0.7060 0.01556 0.00882 -0.0942 0.0191 0.8554
2.750 0.7314 0.01651 0.00990 -0.0934 0.0181 0.8610
3.000 0.7579 0.01790 0.01145 -0.0927 0.0174 0.8673
3.250 0.7827 0.01961 0.01342 -0.0915 0.0169 0.8722
3.500 0.8073 0.02290 0.01718 -0.0898 0.0175 0.8779
4.750 0.8861 0.04652 0.04289 -0.0787 0.0205 0.9074
5.000 0.8980 0.05035 0.04706 -0.0764 0.0204 0.9138
5.250 0.9133 0.05360 0.05066 -0.0743 0.0201 0.9214
5.500 0.9359 0.05633 0.05370 -0.0723 0.0172 0.9294
5.750 0.9442 0.06071 0.05831 -0.0708 0.0163 0.9385
6.000 0.9499 0.06489 0.06269 -0.0694 0.0156 0.9504
6.250 0.9527 0.06901 0.06698 -0.0681 0.0152 0.9999
6.500 0.9586 0.07385 0.07195 -0.0687 0.0148 1.0000
6.750 0.9615 0.07880 0.07702 -0.0696 0.0146 1.0000
7.000 0.9615 0.08383 0.08214 -0.0708 0.0144 1.0000
7.250 0.9581 0.08918 0.08756 -0.0726 0.0142 1.0000
7.500 0.9482 0.09456 0.09299 -0.0746 0.0142 1.0000
7.750 0.9373 0.10160 0.10005 -0.0810 0.0142 1.0000
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Polar data table (+)
Polar graphs
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