GOE 623 AIRFOIL (goe623-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 623 AIRFOIL (goe623-il) Reynolds number: 100,000 Max Cl/Cd: 55.04 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe623-il-100000-n5.txt Download as CSV file: xf-goe623-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 623 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3968 0.09178 0.08678 -0.0447 1.0000 0.0398
-9.500 -0.4014 0.08724 0.08230 -0.0465 1.0000 0.0397
-9.250 -0.4096 0.08257 0.07770 -0.0483 1.0000 0.0396
-9.000 -0.4231 0.07808 0.07332 -0.0495 1.0000 0.0395
-8.750 -0.4446 0.07431 0.06968 -0.0495 1.0000 0.0393
-8.500 -0.4682 0.07003 0.06549 -0.0499 1.0000 0.0391
-8.250 -0.4917 0.06410 0.05954 -0.0513 0.9990 0.0389
-8.000 -0.5001 0.04887 0.04350 -0.0645 0.9858 0.0395
-7.750 -0.4856 0.04295 0.03703 -0.0685 0.9772 0.0408
-7.500 -0.4607 0.03999 0.03380 -0.0708 0.9707 0.0420
-7.250 -0.4354 0.03660 0.02999 -0.0731 0.9640 0.0431
-7.000 -0.4085 0.03395 0.02692 -0.0748 0.9574 0.0451
-6.750 -0.3799 0.03112 0.02349 -0.0764 0.9512 0.0472
-6.500 -0.3481 0.02868 0.02046 -0.0780 0.9462 0.0484
-6.250 -0.3201 0.02701 0.01861 -0.0787 0.9386 0.0499
-6.000 -0.2863 0.02589 0.01738 -0.0804 0.9337 0.0524
-5.750 -0.2578 0.02474 0.01601 -0.0807 0.9258 0.0544
-5.500 -0.2253 0.02350 0.01453 -0.0817 0.9197 0.0562
-5.250 -0.1942 0.02246 0.01328 -0.0823 0.9129 0.0580
-5.000 -0.1645 0.02155 0.01233 -0.0828 0.9052 0.0607
-4.750 -0.1329 0.02082 0.01154 -0.0836 0.8987 0.0645
-4.500 -0.1047 0.02014 0.01074 -0.0835 0.8898 0.0680
-4.250 -0.0744 0.01938 0.00993 -0.0839 0.8826 0.0721
-4.000 -0.0470 0.01879 0.00930 -0.0838 0.8733 0.0784
-3.750 -0.0186 0.01819 0.00874 -0.0838 0.8649 0.0897
-3.500 0.0097 0.01773 0.00840 -0.0838 0.8560 0.1132
-3.250 0.0373 0.01745 0.00810 -0.0836 0.8464 0.1431
-3.000 0.0666 0.01716 0.00772 -0.0836 0.8383 0.1667
-2.750 0.0925 0.01688 0.00750 -0.0832 0.8274 0.1889
-2.500 0.1202 0.01657 0.00726 -0.0831 0.8185 0.2191
-2.250 0.1470 0.01628 0.00708 -0.0828 0.8085 0.2525
-2.000 0.1733 0.01599 0.00690 -0.0825 0.7984 0.2927
-1.750 0.2008 0.01557 0.00667 -0.0822 0.7899 0.3512
-1.500 0.2252 0.01511 0.00658 -0.0816 0.7788 0.4493
-1.250 0.2476 0.01448 0.00654 -0.0801 0.7696 0.6144
-1.000 0.2715 0.01401 0.00657 -0.0779 0.7605 0.7975
-0.750 0.3168 0.01389 0.00649 -0.0805 0.7511 0.9274
-0.500 0.3667 0.01383 0.00624 -0.0848 0.7429 0.9906
-0.250 0.3969 0.01390 0.00617 -0.0854 0.7322 1.0000
0.000 0.4219 0.01397 0.00608 -0.0847 0.7229 1.0000
0.250 0.4471 0.01405 0.00601 -0.0841 0.7135 1.0000
0.500 0.4719 0.01416 0.00600 -0.0835 0.7035 1.0000
0.750 0.4976 0.01424 0.00594 -0.0829 0.6939 1.0000
1.000 0.5228 0.01433 0.00591 -0.0823 0.6830 1.0000
1.250 0.5478 0.01445 0.00592 -0.0816 0.6717 1.0000
1.500 0.5735 0.01456 0.00591 -0.0810 0.6613 1.0000
1.750 0.5993 0.01467 0.00591 -0.0805 0.6507 1.0000
2.000 0.6243 0.01483 0.00599 -0.0798 0.6386 1.0000
2.250 0.6496 0.01498 0.00605 -0.0792 0.6269 1.0000
2.500 0.6752 0.01513 0.00610 -0.0787 0.6156 1.0000
2.750 0.7006 0.01530 0.00620 -0.0781 0.6040 1.0000
3.000 0.7255 0.01549 0.00636 -0.0775 0.5922 1.0000
3.250 0.7507 0.01569 0.00651 -0.0769 0.5811 1.0000
3.500 0.7763 0.01588 0.00663 -0.0764 0.5704 1.0000
3.750 0.8008 0.01609 0.00685 -0.0758 0.5582 1.0000
4.000 0.8256 0.01631 0.00707 -0.0752 0.5463 1.0000
4.250 0.8503 0.01653 0.00727 -0.0746 0.5342 1.0000
4.500 0.8749 0.01675 0.00747 -0.0740 0.5219 1.0000
4.750 0.8990 0.01698 0.00772 -0.0733 0.5083 1.0000
5.000 0.9226 0.01723 0.00798 -0.0725 0.4934 1.0000
5.250 0.9459 0.01748 0.00824 -0.0717 0.4772 1.0000
5.500 0.9686 0.01776 0.00853 -0.0708 0.4596 1.0000
5.750 0.9909 0.01806 0.00885 -0.0699 0.4401 1.0000
6.000 1.0128 0.01840 0.00916 -0.0688 0.4197 1.0000
6.250 1.0338 0.01880 0.00950 -0.0677 0.3988 1.0000
6.500 1.0542 0.01925 0.00991 -0.0665 0.3772 1.0000
6.750 1.0736 0.01978 0.01035 -0.0653 0.3572 1.0000
7.000 1.0926 0.02037 0.01086 -0.0639 0.3396 1.0000
7.250 1.1115 0.02099 0.01144 -0.0627 0.3242 1.0000
7.500 1.1307 0.02162 0.01206 -0.0614 0.3110 1.0000
7.750 1.1502 0.02225 0.01271 -0.0603 0.2999 1.0000
8.000 1.1692 0.02292 0.01337 -0.0591 0.2906 1.0000
8.250 1.1887 0.02358 0.01408 -0.0580 0.2819 1.0000
8.500 1.2079 0.02427 0.01481 -0.0569 0.2743 1.0000
8.750 1.2272 0.02496 0.01556 -0.0558 0.2669 1.0000
9.000 1.2465 0.02569 0.01631 -0.0547 0.2608 1.0000
9.250 1.2656 0.02638 0.01714 -0.0536 0.2541 1.0000
9.500 1.2835 0.02714 0.01792 -0.0524 0.2477 1.0000
9.750 1.3002 0.02787 0.01880 -0.0511 0.2404 1.0000
10.000 1.3146 0.02863 0.01960 -0.0494 0.2336 1.0000
10.250 1.3268 0.02940 0.02052 -0.0474 0.2257 1.0000
10.500 1.3352 0.03027 0.02139 -0.0451 0.2166 1.0000
10.750 1.3410 0.03121 0.02251 -0.0427 0.2056 1.0000
11.000 1.3465 0.03229 0.02370 -0.0406 0.1943 1.0000
11.250 1.3520 0.03351 0.02499 -0.0386 0.1842 1.0000
11.500 1.3574 0.03484 0.02643 -0.0369 0.1743 1.0000
11.750 1.3629 0.03629 0.02803 -0.0354 0.1626 1.0000
12.000 1.3669 0.03795 0.02982 -0.0340 0.1496 1.0000
12.250 1.3692 0.03987 0.03184 -0.0328 0.1335 1.0000
12.500 1.3679 0.04225 0.03426 -0.0316 0.1076 1.0000
12.750 1.3562 0.04575 0.03755 -0.0305 0.0792 1.0000
13.000 1.3426 0.04974 0.04141 -0.0297 0.0654 1.0000
13.250 1.3305 0.05382 0.04550 -0.0292 0.0578 1.0000
13.500 1.3181 0.05816 0.04991 -0.0291 0.0527 1.0000
13.750 1.3052 0.06274 0.05461 -0.0294 0.0492 1.0000
14.000 1.2937 0.06739 0.05943 -0.0300 0.0459 1.0000
14.250 1.2797 0.07258 0.06477 -0.0310 0.0437 1.0000
14.500 1.2633 0.07836 0.07067 -0.0325 0.0421 1.0000
14.750 1.2478 0.08426 0.07670 -0.0342 0.0408 1.0000
15.000 1.2345 0.08999 0.08262 -0.0360 0.0392 1.0000
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Polar data table (+)
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