GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 500,000 Max Cl/Cd: 75.52 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe622-il-500000-n5.txt Download as CSV file: xf-goe622-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 622 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5316 0.08471 0.08256 -0.0159 1.0000 0.0055
-8.750 -0.5390 0.07880 0.07670 -0.0195 1.0000 0.0054
-8.250 -0.6272 0.03371 0.03091 -0.0503 1.0000 0.0048
-8.000 -0.6250 0.02696 0.02348 -0.0500 1.0000 0.0049
-7.750 -0.6122 0.02365 0.01971 -0.0489 1.0000 0.0050
-7.500 -0.5962 0.02129 0.01698 -0.0477 1.0000 0.0052
-7.250 -0.5706 0.01924 0.01456 -0.0482 0.9984 0.0055
-7.000 -0.5382 0.01757 0.01256 -0.0498 0.9954 0.0057
-6.750 -0.5059 0.01634 0.01108 -0.0512 0.9915 0.0060
-6.500 -0.4740 0.01475 0.00926 -0.0526 0.9875 0.0065
-6.250 -0.4422 0.01412 0.00855 -0.0538 0.9823 0.0071
-6.000 -0.4094 0.01347 0.00777 -0.0551 0.9772 0.0079
-5.750 -0.3789 0.01273 0.00689 -0.0558 0.9698 0.0085
-5.500 -0.3475 0.01209 0.00612 -0.0566 0.9640 0.0091
-5.250 -0.3171 0.01124 0.00512 -0.0573 0.9566 0.0104
-5.000 -0.2859 0.01076 0.00459 -0.0581 0.9495 0.0123
-4.750 -0.2548 0.01039 0.00415 -0.0588 0.9417 0.0145
-4.500 -0.2246 0.00992 0.00362 -0.0592 0.9331 0.0198
-4.250 -0.1948 0.00964 0.00328 -0.0596 0.9244 0.0264
-4.000 -0.1665 0.00938 0.00301 -0.0597 0.9148 0.0346
-3.750 -0.1386 0.00914 0.00274 -0.0596 0.9061 0.0423
-3.500 -0.1109 0.00893 0.00250 -0.0595 0.8973 0.0515
-3.250 -0.0836 0.00872 0.00232 -0.0594 0.8880 0.0668
-3.000 -0.0562 0.00857 0.00218 -0.0593 0.8794 0.0848
-2.750 -0.0287 0.00845 0.00203 -0.0591 0.8707 0.0964
-2.500 -0.0011 0.00835 0.00190 -0.0590 0.8617 0.1047
-2.250 0.0264 0.00826 0.00177 -0.0589 0.8523 0.1130
-2.000 0.0539 0.00815 0.00164 -0.0587 0.8424 0.1235
-1.750 0.0814 0.00804 0.00154 -0.0586 0.8322 0.1351
-1.500 0.1087 0.00795 0.00144 -0.0584 0.8194 0.1495
-1.250 0.1355 0.00776 0.00134 -0.0582 0.8031 0.1925
-1.000 0.1620 0.00751 0.00126 -0.0580 0.7855 0.2647
-0.750 0.1885 0.00727 0.00120 -0.0578 0.7649 0.3440
-0.500 0.2143 0.00705 0.00114 -0.0574 0.7371 0.4342
-0.250 0.2390 0.00689 0.00109 -0.0568 0.6919 0.5311
0.000 0.2629 0.00675 0.00108 -0.0560 0.6462 0.6410
0.250 0.2863 0.00656 0.00112 -0.0550 0.6160 0.7544
0.500 0.3085 0.00638 0.00118 -0.0535 0.5974 0.8564
0.750 0.3355 0.00634 0.00122 -0.0529 0.5821 0.9373
1.000 0.3776 0.00642 0.00124 -0.0560 0.5646 0.9930
1.250 0.4076 0.00653 0.00127 -0.0565 0.5496 1.0000
1.500 0.4342 0.00664 0.00132 -0.0563 0.5367 1.0000
1.750 0.4609 0.00677 0.00138 -0.0560 0.5224 1.0000
2.000 0.4875 0.00690 0.00144 -0.0558 0.5055 1.0000
2.250 0.5141 0.00704 0.00152 -0.0555 0.4867 1.0000
2.500 0.5402 0.00725 0.00161 -0.0552 0.4567 1.0000
2.750 0.5660 0.00750 0.00171 -0.0549 0.4179 1.0000
3.000 0.5913 0.00783 0.00184 -0.0545 0.3709 1.0000
3.250 0.6165 0.00819 0.00203 -0.0541 0.3280 1.0000
3.500 0.6420 0.00853 0.00222 -0.0538 0.2933 1.0000
3.750 0.6676 0.00886 0.00243 -0.0535 0.2617 1.0000
4.000 0.6924 0.00932 0.00268 -0.0531 0.2171 1.0000
4.250 0.7170 0.00981 0.00296 -0.0528 0.1689 1.0000
4.500 0.7411 0.01036 0.00328 -0.0523 0.1257 1.0000
4.750 0.7667 0.01070 0.00356 -0.0520 0.1090 1.0000
5.000 0.7923 0.01103 0.00383 -0.0517 0.0918 1.0000
5.250 0.8159 0.01165 0.00424 -0.0512 0.0511 1.0000
5.500 0.8391 0.01235 0.00476 -0.0506 0.0187 1.0000
5.750 0.8642 0.01276 0.00519 -0.0501 0.0133 1.0000
6.000 0.8894 0.01317 0.00568 -0.0497 0.0111 1.0000
6.250 0.9144 0.01357 0.00617 -0.0492 0.0097 1.0000
6.500 0.9388 0.01407 0.00677 -0.0487 0.0087 1.0000
6.750 0.9621 0.01471 0.00750 -0.0480 0.0078 1.0000
7.000 0.9840 0.01557 0.00848 -0.0471 0.0072 1.0000
7.250 1.0075 0.01613 0.00912 -0.0465 0.0067 1.0000
7.500 1.0306 0.01672 0.00978 -0.0458 0.0061 1.0000
7.750 1.0529 0.01740 0.01052 -0.0451 0.0056 1.0000
8.000 1.0744 0.01817 0.01136 -0.0442 0.0052 1.0000
8.250 1.0941 0.01913 0.01244 -0.0431 0.0050 1.0000
8.500 1.1105 0.02054 0.01396 -0.0416 0.0047 1.0000
8.750 1.1297 0.02155 0.01509 -0.0404 0.0045 1.0000
9.000 1.1475 0.02272 0.01640 -0.0390 0.0044 1.0000
9.250 1.1641 0.02404 0.01788 -0.0375 0.0042 1.0000
9.500 1.1795 0.02551 0.01951 -0.0359 0.0041 1.0000
9.750 1.1944 0.02698 0.02115 -0.0343 0.0039 1.0000
10.000 1.2089 0.02833 0.02265 -0.0327 0.0037 1.0000
10.250 1.2227 0.02960 0.02406 -0.0312 0.0035 1.0000
10.500 1.2353 0.03080 0.02538 -0.0296 0.0033 1.0000
10.750 1.2449 0.03205 0.02675 -0.0276 0.0032 1.0000
11.000 1.2501 0.03351 0.02832 -0.0252 0.0031 1.0000
11.250 1.2507 0.03553 0.03050 -0.0226 0.0031 1.0000
11.500 1.2461 0.03817 0.03334 -0.0200 0.0030 1.0000
11.750 1.2392 0.04119 0.03664 -0.0179 0.0030 1.0000
12.000 1.2336 0.04420 0.03988 -0.0166 0.0029 1.0000
12.250 1.2258 0.04769 0.04359 -0.0160 0.0029 1.0000
12.500 1.2159 0.05177 0.04789 -0.0162 0.0029 1.0000
12.750 1.2035 0.05658 0.05293 -0.0173 0.0029 1.0000
13.000 1.1902 0.06203 0.05858 -0.0195 0.0028 1.0000
13.250 1.1747 0.06829 0.06505 -0.0226 0.0028 1.0000
13.500 1.1585 0.07516 0.07209 -0.0264 0.0029 1.0000
13.750 1.1406 0.08275 0.07985 -0.0307 0.0029 1.0000
14.000 1.1228 0.09079 0.08804 -0.0355 0.0029 1.0000
14.250 1.1045 0.09944 0.09682 -0.0405 0.0029 1.0000
14.500 1.0860 0.10858 0.10606 -0.0457 0.0029 1.0000
14.750 1.0672 0.11809 0.11564 -0.0509 0.0029 1.0000
15.000 1.0488 0.12761 0.12524 -0.0559 0.0030 1.0000
15.250 0.8170 0.12404 0.12189 -0.0350 0.0034 1.0000
15.500 0.7632 0.13758 0.13559 -0.0414 0.0038 1.0000
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Polar data table (+)
Polar graphs
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