GOE 602 AIRFOIL (goe602-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 602 AIRFOIL (goe602-il) Reynolds number: 50,000 Max Cl/Cd: 37.84 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe602-il-50000.txt Download as CSV file: xf-goe602-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3826 0.10191 0.09539 -0.0240 1.0000 0.2582
-7.750 -0.3683 0.09778 0.09127 -0.0222 1.0000 0.2701
-7.500 -0.3680 0.09498 0.08856 -0.0205 1.0000 0.2799
-7.250 -0.4000 0.09540 0.08919 -0.0178 1.0000 0.2879
-7.000 -0.3917 0.09197 0.08580 -0.0156 1.0000 0.3015
-6.750 -0.3918 0.08931 0.08321 -0.0130 1.0000 0.3150
-6.500 -0.3978 0.08723 0.08123 -0.0098 1.0000 0.3294
-6.250 -0.4058 0.08529 0.07940 -0.0067 1.0000 0.3446
-6.000 -0.4189 0.08370 0.07792 -0.0037 1.0000 0.3612
-5.750 -0.4099 0.08048 0.07475 -0.0001 1.0000 0.3835
-5.500 -0.4194 0.07878 0.07315 0.0033 1.0000 0.4046
-5.250 -0.4210 0.07627 0.07069 0.0068 1.0000 0.4257
-5.000 -0.4351 0.07503 0.06955 0.0106 1.0000 0.4515
-4.000 -0.3486 0.04225 0.03439 -0.0414 1.0000 0.1528
-3.750 -0.3251 0.03904 0.03068 -0.0420 1.0000 0.1513
-3.500 -0.2998 0.03614 0.02718 -0.0424 1.0000 0.1492
-3.250 -0.2739 0.03363 0.02408 -0.0424 1.0000 0.1472
-3.000 -0.2499 0.03184 0.02190 -0.0420 1.0000 0.1499
-2.750 -0.2259 0.03050 0.02006 -0.0415 1.0000 0.1564
-2.500 -0.2031 0.02905 0.01846 -0.0408 1.0000 0.1616
-2.250 -0.1796 0.02799 0.01716 -0.0401 1.0000 0.1684
-2.000 -0.1578 0.02708 0.01621 -0.0391 1.0000 0.1802
-1.750 -0.1353 0.02641 0.01544 -0.0383 1.0000 0.2007
-1.500 -0.1111 0.02564 0.01487 -0.0378 1.0000 0.2330
-1.250 -0.0835 0.02155 0.01408 -0.0354 1.0000 1.0000
-1.000 -0.0628 0.02200 0.01380 -0.0350 1.0000 1.0000
-0.750 -0.0436 0.02249 0.01386 -0.0345 1.0000 1.0000
-0.500 -0.0251 0.02304 0.01407 -0.0341 1.0000 1.0000
-0.250 -0.0070 0.02363 0.01438 -0.0338 1.0000 1.0000
0.000 0.0108 0.02429 0.01476 -0.0335 1.0000 1.0000
0.250 0.0281 0.02499 0.01525 -0.0332 1.0000 1.0000
0.500 0.0691 0.02625 0.01625 -0.0374 0.9893 1.0000
0.750 0.1188 0.02764 0.01738 -0.0430 0.9737 1.0000
1.000 0.1666 0.02890 0.01843 -0.0481 0.9581 1.0000
1.250 0.2101 0.02997 0.01935 -0.0523 0.9422 1.0000
1.500 0.2483 0.03087 0.02014 -0.0553 0.9256 1.0000
1.750 0.2862 0.03176 0.02093 -0.0582 0.9093 1.0000
2.000 0.3235 0.03260 0.02171 -0.0608 0.8932 1.0000
2.250 0.3606 0.03340 0.02247 -0.0633 0.8772 1.0000
2.500 0.3975 0.03415 0.02320 -0.0655 0.8613 1.0000
2.750 0.4341 0.03484 0.02389 -0.0676 0.8458 1.0000
3.000 0.4709 0.03547 0.02454 -0.0695 0.8303 1.0000
3.250 0.5080 0.03600 0.02512 -0.0713 0.8151 1.0000
3.500 0.5452 0.03647 0.02564 -0.0730 0.8000 1.0000
3.750 0.5747 0.03705 0.02630 -0.0734 0.7841 1.0000
4.000 0.6011 0.03768 0.02700 -0.0735 0.7674 1.0000
4.250 0.6311 0.03819 0.02761 -0.0738 0.7515 1.0000
4.500 0.6629 0.03860 0.02816 -0.0742 0.7358 1.0000
4.750 0.6958 0.03891 0.02860 -0.0746 0.7203 1.0000
5.000 0.7298 0.03910 0.02895 -0.0750 0.7051 1.0000
5.250 0.7651 0.03916 0.02922 -0.0753 0.6901 1.0000
5.500 0.8028 0.03905 0.02931 -0.0757 0.6749 1.0000
5.750 0.8422 0.03877 0.02925 -0.0761 0.6595 1.0000
6.000 0.8823 0.03835 0.02910 -0.0763 0.6438 1.0000
6.250 0.9331 0.03709 0.02813 -0.0770 0.6273 1.0000
6.500 0.9569 0.03722 0.02845 -0.0751 0.6058 1.0000
6.750 1.0170 0.03454 0.02606 -0.0754 0.5806 1.0000
7.000 1.0604 0.03310 0.02478 -0.0747 0.5538 1.0000
7.250 1.0847 0.03191 0.02366 -0.0713 0.5178 1.0000
7.500 1.1135 0.03040 0.02211 -0.0685 0.4800 1.0000
7.750 1.1269 0.03024 0.02202 -0.0648 0.4471 1.0000
8.000 1.1380 0.03011 0.02195 -0.0609 0.4131 1.0000
8.250 1.1391 0.03010 0.02194 -0.0557 0.3731 1.0000
8.500 1.1337 0.03055 0.02233 -0.0499 0.3273 1.0000
8.750 1.1205 0.03165 0.02315 -0.0435 0.2688 1.0000
9.000 1.1010 0.03397 0.02483 -0.0373 0.2025 1.0000
9.250 1.0872 0.03705 0.02726 -0.0327 0.1558 1.0000
9.500 1.0842 0.03987 0.02972 -0.0296 0.1293 1.0000
9.750 1.0946 0.04240 0.03202 -0.0275 0.1127 1.0000
10.000 1.1333 0.04497 0.03456 -0.0274 0.0989 1.0000
10.250 1.1840 0.04848 0.03801 -0.0296 0.0875 1.0000
10.500 1.2042 0.05152 0.04158 -0.0285 0.0837 1.0000
10.750 1.2250 0.05527 0.04577 -0.0278 0.0817 1.0000
11.000 1.2337 0.05903 0.04994 -0.0260 0.0807 1.0000
11.250 1.2328 0.06276 0.05406 -0.0234 0.0805 1.0000
11.500 1.2241 0.06647 0.05814 -0.0206 0.0806 1.0000
11.750 1.2094 0.07029 0.06229 -0.0179 0.0810 1.0000
12.000 1.1900 0.07439 0.06670 -0.0158 0.0814 1.0000
12.250 1.1670 0.07893 0.07152 -0.0146 0.0821 1.0000
12.500 1.1420 0.08406 0.07689 -0.0146 0.0830 1.0000
12.750 1.1152 0.08992 0.08295 -0.0158 0.0839 1.0000
13.000 1.0880 0.09658 0.08977 -0.0182 0.0849 1.0000
13.250 1.0623 0.10396 0.09725 -0.0217 0.0859 1.0000
13.500 1.0412 0.11171 0.10507 -0.0255 0.0870 1.0000
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Polar data table (+)
Polar graphs
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