GOE 373 AIRFOIL (goe373-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 373 AIRFOIL (goe373-il) Reynolds number: 100,000 Max Cl/Cd: 61.1 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe373-il-100000-n5.txt Download as CSV file: xf-goe373-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 373 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2893 0.09385 0.08917 -0.0232 1.0000 0.0321
-7.250 -0.2932 0.09244 0.08786 -0.0235 1.0000 0.0333
-7.000 -0.2933 0.09116 0.08668 -0.0268 1.0000 0.0338
-6.750 -0.2891 0.08956 0.08513 -0.0311 1.0000 0.0341
-6.500 -0.2871 0.08588 0.08154 -0.0289 1.0000 0.0344
-6.250 -0.2862 0.08270 0.07841 -0.0257 1.0000 0.0351
-6.000 -0.2857 0.08035 0.07611 -0.0244 1.0000 0.0358
-5.750 -0.2651 0.07677 0.07252 -0.0285 0.9949 0.0373
-5.500 -0.2288 0.07256 0.06823 -0.0375 0.9860 0.0404
-5.250 -0.1875 0.06787 0.06343 -0.0484 0.9761 0.0430
-5.000 -0.1646 0.06418 0.05973 -0.0507 0.9675 0.0450
-4.750 -0.1285 0.06043 0.05587 -0.0573 0.9587 0.0488
-4.500 -0.0811 0.05621 0.05142 -0.0673 0.9497 0.0526
-4.250 -0.0587 0.05310 0.04831 -0.0686 0.9402 0.0563
-4.000 -0.0102 0.04949 0.04444 -0.0771 0.9329 0.0638
-3.750 0.0144 0.04661 0.04154 -0.0786 0.9217 0.0675
-3.500 0.0610 0.04398 0.03850 -0.0851 0.9113 0.0755
-3.250 0.0877 0.04079 0.03536 -0.0867 0.9026 0.0794
-3.000 0.1269 0.03861 0.03282 -0.0906 0.8905 0.0896
-2.750 0.1548 0.03628 0.03047 -0.0920 0.8782 0.0988
-2.500 0.1864 0.03396 0.02798 -0.0940 0.8642 0.1109
-2.250 0.2185 0.03192 0.02574 -0.0957 0.8483 0.1250
-2.000 0.2492 0.03012 0.02374 -0.0969 0.8316 0.1410
-1.750 0.2787 0.02861 0.02198 -0.0978 0.8152 0.1666
-1.250 0.3564 0.02501 0.01713 -0.0982 0.7866 0.0734
-1.000 0.3841 0.02328 0.01525 -0.0982 0.7720 0.0696
-0.750 0.4158 0.02253 0.01389 -0.0976 0.7573 0.0620
-0.500 0.4426 0.02108 0.01236 -0.0975 0.7423 0.0593
-0.250 0.4706 0.02013 0.01114 -0.0970 0.7272 0.0563
0.000 0.4987 0.01939 0.01007 -0.0963 0.7119 0.0531
0.250 0.5261 0.01900 0.00941 -0.0956 0.6963 0.0514
0.500 0.5527 0.01832 0.00862 -0.0951 0.6807 0.0506
0.750 0.5791 0.01778 0.00797 -0.0945 0.6653 0.0499
1.000 0.6054 0.01733 0.00743 -0.0939 0.6498 0.0494
1.250 0.6313 0.01696 0.00698 -0.0932 0.6342 0.0492
1.500 0.6575 0.01671 0.00664 -0.0926 0.6188 0.0508
1.750 0.6826 0.01645 0.00634 -0.0919 0.6039 0.0521
2.000 0.7082 0.01627 0.00607 -0.0912 0.5899 0.0525
2.250 0.7346 0.01619 0.00589 -0.0907 0.5764 0.0528
2.500 0.7609 0.01619 0.00576 -0.0902 0.5632 0.0533
2.750 0.7869 0.01626 0.00571 -0.0896 0.5496 0.0542
3.000 0.8128 0.01638 0.00572 -0.0891 0.5360 0.0554
3.250 0.8385 0.01653 0.00579 -0.0885 0.5225 0.0572
3.500 0.8640 0.01671 0.00591 -0.0879 0.5083 0.0601
3.750 0.8894 0.01688 0.00608 -0.0874 0.4939 0.0744
4.000 0.9174 0.01563 0.00640 -0.0877 0.4799 1.0000
4.250 0.9423 0.01593 0.00662 -0.0870 0.4668 1.0000
4.500 0.9670 0.01623 0.00686 -0.0864 0.4539 1.0000
4.750 0.9915 0.01653 0.00713 -0.0857 0.4407 1.0000
5.000 1.0159 0.01685 0.00741 -0.0850 0.4284 1.0000
5.250 1.0402 0.01717 0.00773 -0.0844 0.4166 1.0000
5.500 1.0646 0.01750 0.00810 -0.0837 0.4053 1.0000
5.750 1.0888 0.01785 0.00850 -0.0831 0.3949 1.0000
6.000 1.1121 0.01820 0.00886 -0.0823 0.3813 1.0000
6.250 1.1339 0.01857 0.00920 -0.0813 0.3609 1.0000
6.500 1.1553 0.01900 0.00957 -0.0802 0.3410 1.0000
6.750 1.1765 0.01946 0.01004 -0.0792 0.3209 1.0000
7.000 1.1964 0.01998 0.01051 -0.0780 0.2942 1.0000
7.250 1.2167 0.02051 0.01103 -0.0769 0.2658 1.0000
7.500 1.2354 0.02118 0.01158 -0.0757 0.2263 1.0000
7.750 1.2519 0.02210 0.01229 -0.0742 0.1896 1.0000
8.000 1.2687 0.02304 0.01313 -0.0728 0.1675 1.0000
8.250 1.2847 0.02408 0.01409 -0.0714 0.1465 1.0000
8.500 1.2996 0.02523 0.01512 -0.0698 0.1101 1.0000
8.750 1.3091 0.02684 0.01640 -0.0678 0.0648 1.0000
9.000 1.3133 0.02886 0.01815 -0.0651 0.0275 1.0000
9.250 1.3220 0.03036 0.01966 -0.0629 0.0244 1.0000
9.500 1.3305 0.03177 0.02118 -0.0605 0.0227 1.0000
9.750 1.3366 0.03316 0.02273 -0.0579 0.0216 1.0000
10.000 1.3406 0.03471 0.02447 -0.0553 0.0205 1.0000
10.250 1.3422 0.03649 0.02644 -0.0528 0.0196 1.0000
10.500 1.3415 0.03854 0.02869 -0.0504 0.0188 1.0000
10.750 1.3392 0.04085 0.03122 -0.0485 0.0183 1.0000
11.000 1.3355 0.04347 0.03404 -0.0470 0.0180 1.0000
11.250 1.3305 0.04642 0.03719 -0.0460 0.0178 1.0000
11.500 1.3239 0.04979 0.04075 -0.0455 0.0175 1.0000
11.750 1.3165 0.05350 0.04465 -0.0455 0.0174 1.0000
12.000 1.3082 0.05757 0.04890 -0.0459 0.0172 1.0000
12.250 1.3000 0.06186 0.05336 -0.0467 0.0171 1.0000
12.500 1.2921 0.06630 0.05796 -0.0477 0.0170 1.0000
12.750 1.2844 0.07084 0.06265 -0.0488 0.0169 1.0000
13.000 1.2773 0.07533 0.06728 -0.0498 0.0168 1.0000
13.250 1.2721 0.07953 0.07161 -0.0506 0.0166 1.0000
13.500 1.2699 0.08336 0.07554 -0.0513 0.0166 1.0000
13.750 1.2694 0.08701 0.07932 -0.0519 0.0165 1.0000
14.000 1.2695 0.09060 0.08307 -0.0526 0.0163 1.0000
14.250 1.2694 0.09437 0.08699 -0.0534 0.0161 1.0000
14.500 1.2686 0.09835 0.09114 -0.0545 0.0158 1.0000
14.750 1.2671 0.10260 0.09556 -0.0558 0.0155 1.0000
15.000 1.2646 0.10710 0.10024 -0.0574 0.0152 1.0000
15.250 1.2624 0.11163 0.10494 -0.0591 0.0150 1.0000
15.500 1.2591 0.11651 0.10999 -0.0611 0.0147 1.0000
15.750 1.2553 0.12160 0.11527 -0.0634 0.0146 1.0000
16.000 1.2506 0.12702 0.12087 -0.0660 0.0146 1.0000
16.250 1.2447 0.13287 0.12691 -0.0691 0.0146 1.0000
16.500 1.2376 0.13918 0.13340 -0.0727 0.0146 1.0000
16.750 1.2296 0.14594 0.14034 -0.0767 0.0146 1.0000
17.000 1.2207 0.15318 0.14776 -0.0812 0.0147 1.0000
17.250 1.2109 0.16100 0.15575 -0.0862 0.0148 1.0000
17.500 1.2001 0.16944 0.16435 -0.0916 0.0149 1.0000
17.750 1.1886 0.17865 0.17370 -0.0976 0.0150 1.0000
18.000 1.1769 0.18853 0.18369 -0.1038 0.0153 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 373 AIRFOIL (goe373-il)