GOE 623 AIRFOIL (goe623-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 623 AIRFOIL (goe623-il) Reynolds number: 50,000 Max Cl/Cd: 36 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe623-il-50000-n5.txt Download as CSV file: xf-goe623-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3785 0.10474 0.09763 -0.0418 1.0000 0.0696
-9.500 -0.3742 0.10121 0.09413 -0.0420 1.0000 0.0690
-9.250 -0.3740 0.09756 0.09055 -0.0428 1.0000 0.0685
-9.000 -0.3783 0.09378 0.08687 -0.0440 1.0000 0.0684
-8.750 -0.3865 0.09010 0.08330 -0.0449 1.0000 0.0684
-8.500 -0.3990 0.08672 0.08005 -0.0451 1.0000 0.0684
-8.250 -0.4133 0.08332 0.07678 -0.0454 1.0000 0.0683
-8.000 -0.4270 0.07963 0.07318 -0.0458 1.0000 0.0681
-7.750 -0.4410 0.07590 0.06950 -0.0457 1.0000 0.0680
-7.500 -0.4542 0.07203 0.06564 -0.0455 1.0000 0.0679
-7.250 -0.4654 0.06797 0.06152 -0.0453 1.0000 0.0678
-7.000 -0.4736 0.06372 0.05716 -0.0452 1.0000 0.0677
-6.750 -0.4779 0.05933 0.05256 -0.0453 1.0000 0.0676
-6.500 -0.4771 0.05483 0.04776 -0.0456 0.9997 0.0677
-6.250 -0.4518 0.04803 0.04013 -0.0512 0.9918 0.0695
-6.000 -0.4208 0.04492 0.03676 -0.0545 0.9848 0.0722
-5.750 -0.3899 0.04190 0.03336 -0.0572 0.9770 0.0741
-5.500 -0.3576 0.03875 0.02969 -0.0598 0.9695 0.0760
-5.250 -0.3222 0.03608 0.02643 -0.0624 0.9626 0.0798
-5.000 -0.2895 0.03382 0.02361 -0.0641 0.9543 0.0839
-4.750 -0.2517 0.03231 0.02195 -0.0666 0.9480 0.0874
-4.500 -0.2199 0.03093 0.02031 -0.0677 0.9388 0.0915
-4.250 -0.1801 0.02959 0.01868 -0.0700 0.9329 0.0982
-4.000 -0.1490 0.02885 0.01785 -0.0709 0.9230 0.1072
-3.750 -0.1083 0.02790 0.01679 -0.0733 0.9171 0.1198
-3.500 -0.0769 0.02714 0.01591 -0.0740 0.9074 0.1374
-3.250 -0.0365 0.02633 0.01509 -0.0765 0.9014 0.1703
-3.000 -0.0076 0.02584 0.01469 -0.0771 0.8909 0.2077
-2.750 0.0294 0.02533 0.01430 -0.0791 0.8843 0.2593
-2.500 0.0572 0.02493 0.01397 -0.0792 0.8737 0.2950
-2.250 0.0887 0.02441 0.01358 -0.0800 0.8653 0.3413
-2.000 0.1179 0.02378 0.01330 -0.0803 0.8564 0.4098
-1.750 0.1414 0.02303 0.01319 -0.0792 0.8470 0.5340
-1.500 0.1716 0.02217 0.01320 -0.0773 0.8402 0.8053
-1.250 0.2376 0.02199 0.01277 -0.0842 0.8337 1.0000
-1.000 0.2655 0.02212 0.01261 -0.0843 0.8241 1.0000
-0.750 0.2897 0.02235 0.01259 -0.0838 0.8135 1.0000
-0.500 0.3216 0.02240 0.01238 -0.0844 0.8061 1.0000
-0.250 0.3431 0.02272 0.01253 -0.0835 0.7949 1.0000
0.000 0.3705 0.02289 0.01252 -0.0834 0.7863 1.0000
0.250 0.3965 0.02311 0.01257 -0.0830 0.7770 1.0000
0.500 0.4203 0.02341 0.01273 -0.0824 0.7672 1.0000
0.750 0.4497 0.02351 0.01269 -0.0825 0.7596 1.0000
1.000 0.4716 0.02388 0.01297 -0.0816 0.7489 1.0000
1.250 0.5014 0.02394 0.01291 -0.0816 0.7414 1.0000
1.500 0.5243 0.02424 0.01314 -0.0808 0.7306 1.0000
1.750 0.5487 0.02448 0.01331 -0.0800 0.7203 1.0000
2.000 0.5792 0.02443 0.01316 -0.0800 0.7121 1.0000
2.250 0.6013 0.02474 0.01344 -0.0789 0.7002 1.0000
2.500 0.6262 0.02492 0.01357 -0.0782 0.6895 1.0000
2.750 0.6573 0.02478 0.01336 -0.0780 0.6808 1.0000
3.000 0.6799 0.02504 0.01361 -0.0769 0.6682 1.0000
3.250 0.7041 0.02521 0.01376 -0.0760 0.6560 1.0000
3.500 0.7302 0.02528 0.01381 -0.0752 0.6440 1.0000
3.750 0.7586 0.02523 0.01371 -0.0747 0.6326 1.0000
4.000 0.7838 0.02534 0.01381 -0.0738 0.6196 1.0000
4.250 0.8070 0.02557 0.01406 -0.0727 0.6057 1.0000
4.500 0.8305 0.02578 0.01430 -0.0717 0.5917 1.0000
4.750 0.8541 0.02600 0.01453 -0.0707 0.5775 1.0000
5.000 0.8775 0.02623 0.01480 -0.0696 0.5630 1.0000
5.250 0.9005 0.02648 0.01507 -0.0685 0.5479 1.0000
5.500 0.9228 0.02676 0.01538 -0.0674 0.5322 1.0000
5.750 0.9443 0.02706 0.01575 -0.0661 0.5159 1.0000
6.000 0.9653 0.02741 0.01614 -0.0648 0.4990 1.0000
6.250 0.9858 0.02778 0.01655 -0.0635 0.4817 1.0000
6.500 1.0061 0.02816 0.01697 -0.0621 0.4640 1.0000
6.750 1.0264 0.02857 0.01742 -0.0608 0.4467 1.0000
7.000 1.0458 0.02905 0.01791 -0.0594 0.4292 1.0000
7.250 1.0640 0.02965 0.01855 -0.0580 0.4116 1.0000
7.500 1.0828 0.03026 0.01918 -0.0566 0.3955 1.0000
7.750 1.1022 0.03092 0.01986 -0.0554 0.3811 1.0000
8.000 1.1227 0.03159 0.02052 -0.0543 0.3683 1.0000
8.250 1.1420 0.03240 0.02137 -0.0532 0.3560 1.0000
8.500 1.1603 0.03332 0.02237 -0.0520 0.3445 1.0000
8.750 1.1803 0.03420 0.02333 -0.0511 0.3345 1.0000
9.000 1.2015 0.03508 0.02427 -0.0503 0.3255 1.0000
9.250 1.2198 0.03622 0.02557 -0.0493 0.3174 1.0000
9.500 1.2424 0.03716 0.02659 -0.0487 0.3102 1.0000
9.750 1.2590 0.03843 0.02808 -0.0476 0.3030 1.0000
10.000 1.2786 0.03954 0.02934 -0.0468 0.2961 1.0000
10.250 1.2975 0.04078 0.03076 -0.0460 0.2901 1.0000
10.500 1.3105 0.04229 0.03254 -0.0446 0.2840 1.0000
10.750 1.3324 0.04320 0.03352 -0.0439 0.2767 1.0000
11.000 1.3312 0.04492 0.03556 -0.0410 0.2689 1.0000
11.250 1.3405 0.04591 0.03662 -0.0390 0.2596 1.0000
11.500 1.3393 0.04734 0.03821 -0.0362 0.2503 1.0000
11.750 1.3352 0.04910 0.04015 -0.0336 0.2414 1.0000
12.000 1.3372 0.05025 0.04132 -0.0315 0.2307 1.0000
12.250 1.3239 0.05290 0.04421 -0.0292 0.2219 1.0000
12.500 1.3157 0.05525 0.04669 -0.0276 0.2123 1.0000
12.750 1.3101 0.05734 0.04879 -0.0263 0.2016 1.0000
13.000 1.2922 0.06146 0.05321 -0.0257 0.1933 1.0000
13.250 1.2811 0.06488 0.05673 -0.0254 0.1835 1.0000
13.500 1.2721 0.06814 0.06005 -0.0254 0.1729 1.0000
13.750 1.2520 0.07379 0.06603 -0.0266 0.1645 1.0000
14.000 1.2380 0.07854 0.07092 -0.0276 0.1545 1.0000
14.250 1.2249 0.08337 0.07588 -0.0288 0.1437 1.0000
14.500 1.2031 0.09032 0.08309 -0.0312 0.1346 1.0000
14.750 1.1836 0.09707 0.08997 -0.0337 0.1249 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 623 AIRFOIL (goe623-il)