GOE 507 AIRFOIL (goe507-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 507 AIRFOIL (goe507-il) Reynolds number: 100,000 Max Cl/Cd: 53.56 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe507-il-100000-n5.txt Download as CSV file: xf-goe507-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 507 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4437 0.09158 0.08635 -0.0352 1.0000 0.0385
-9.250 -0.4515 0.08657 0.08142 -0.0377 1.0000 0.0391
-9.000 -0.4608 0.08172 0.07666 -0.0400 1.0000 0.0395
-8.750 -0.4755 0.07679 0.07183 -0.0422 1.0000 0.0399
-8.500 -0.4940 0.07102 0.06618 -0.0452 1.0000 0.0402
-8.250 -0.5118 0.06386 0.05903 -0.0488 1.0000 0.0404
-8.000 -0.5210 0.05944 0.05460 -0.0492 1.0000 0.0411
-7.750 -0.5185 0.05841 0.05359 -0.0473 1.0000 0.0421
-7.500 -0.5178 0.05654 0.05169 -0.0458 1.0000 0.0432
-7.250 -0.5215 0.05288 0.04792 -0.0450 1.0000 0.0443
-7.000 -0.5239 0.04866 0.04350 -0.0443 1.0000 0.0463
-6.750 -0.5239 0.04188 0.03618 -0.0453 0.9986 0.0493
-6.500 -0.5044 0.03444 0.02755 -0.0489 0.9922 0.0540
-6.250 -0.4762 0.03350 0.02663 -0.0502 0.9879 0.0573
-6.000 -0.4488 0.03126 0.02393 -0.0516 0.9834 0.0618
-5.750 -0.4213 0.02866 0.02053 -0.0527 0.9785 0.0661
-5.500 -0.3923 0.02740 0.01923 -0.0538 0.9746 0.0696
-5.250 -0.3642 0.02617 0.01774 -0.0545 0.9699 0.0730
-5.000 -0.3323 0.02489 0.01602 -0.0557 0.9665 0.0765
-4.750 -0.3042 0.02376 0.01452 -0.0560 0.9615 0.0789
-4.500 -0.2736 0.02262 0.01326 -0.0569 0.9576 0.0806
-4.250 -0.2400 0.02174 0.01227 -0.0583 0.9548 0.0825
-4.000 -0.2133 0.02114 0.01156 -0.0583 0.9490 0.0854
-3.750 -0.1805 0.02049 0.01077 -0.0594 0.9452 0.0882
-3.500 -0.1473 0.01984 0.00999 -0.0605 0.9416 0.0897
-3.250 -0.1171 0.01928 0.00931 -0.0610 0.9349 0.0913
-3.000 -0.0790 0.01859 0.00862 -0.0630 0.9306 0.0936
-2.750 -0.0496 0.01811 0.00814 -0.0633 0.9228 0.0964
-2.500 -0.0146 0.01768 0.00768 -0.0646 0.9177 0.1000
-2.250 0.0148 0.01736 0.00733 -0.0648 0.9116 0.1044
-2.000 0.0445 0.01704 0.00704 -0.0651 0.9054 0.1102
-1.750 0.0789 0.01669 0.00676 -0.0663 0.9012 0.1228
-1.500 0.1040 0.01645 0.00669 -0.0658 0.8933 0.1521
-1.250 0.1360 0.01617 0.00654 -0.0665 0.8878 0.1993
-1.000 0.1657 0.01593 0.00640 -0.0668 0.8815 0.2367
-0.750 0.1943 0.01567 0.00629 -0.0668 0.8742 0.2781
-0.500 0.2277 0.01514 0.00615 -0.0679 0.8694 0.3684
-0.250 0.2465 0.01435 0.00618 -0.0660 0.8599 0.5864
0.250 0.3709 0.01343 0.00594 -0.0787 0.8525 1.0000
0.500 0.3955 0.01346 0.00590 -0.0778 0.8423 1.0000
0.750 0.4201 0.01349 0.00588 -0.0769 0.8318 1.0000
1.000 0.4482 0.01344 0.00577 -0.0765 0.8227 1.0000
1.250 0.4717 0.01348 0.00577 -0.0753 0.8105 1.0000
1.500 0.4958 0.01350 0.00576 -0.0742 0.7976 1.0000
1.750 0.5204 0.01348 0.00571 -0.0731 0.7837 1.0000
2.000 0.5450 0.01346 0.00566 -0.0719 0.7681 1.0000
2.250 0.5694 0.01344 0.00562 -0.0707 0.7514 1.0000
2.500 0.5940 0.01346 0.00562 -0.0696 0.7346 1.0000
2.750 0.6186 0.01349 0.00565 -0.0686 0.7175 1.0000
3.000 0.6434 0.01354 0.00568 -0.0676 0.6997 1.0000
3.250 0.6667 0.01363 0.00578 -0.0663 0.6776 1.0000
3.500 0.6907 0.01370 0.00579 -0.0650 0.6523 1.0000
3.750 0.7136 0.01381 0.00583 -0.0636 0.6207 1.0000
4.000 0.7361 0.01399 0.00588 -0.0621 0.5843 1.0000
4.250 0.7583 0.01426 0.00599 -0.0606 0.5491 1.0000
4.500 0.7801 0.01461 0.00619 -0.0592 0.5167 1.0000
4.750 0.8018 0.01498 0.00650 -0.0578 0.4877 1.0000
5.000 0.8232 0.01537 0.00683 -0.0565 0.4581 1.0000
5.250 0.8443 0.01577 0.00719 -0.0551 0.4286 1.0000
5.500 0.8643 0.01623 0.00758 -0.0536 0.3928 1.0000
5.750 0.8833 0.01677 0.00798 -0.0520 0.3535 1.0000
6.000 0.9021 0.01735 0.00845 -0.0504 0.3142 1.0000
6.250 0.9195 0.01806 0.00899 -0.0487 0.2693 1.0000
6.500 0.9349 0.01898 0.00966 -0.0468 0.2142 1.0000
6.750 0.9490 0.02011 0.01045 -0.0448 0.1502 1.0000
7.000 0.9632 0.02129 0.01134 -0.0429 0.1149 1.0000
7.250 0.9781 0.02236 0.01230 -0.0410 0.0988 1.0000
7.500 0.9935 0.02335 0.01332 -0.0392 0.0885 1.0000
7.750 1.0074 0.02444 0.01442 -0.0373 0.0804 1.0000
8.000 1.0210 0.02553 0.01556 -0.0353 0.0729 1.0000
8.250 1.0343 0.02665 0.01677 -0.0333 0.0662 1.0000
8.500 1.0455 0.02792 0.01805 -0.0311 0.0604 1.0000
8.750 1.0595 0.02907 0.01937 -0.0292 0.0546 1.0000
9.000 1.0710 0.03029 0.02065 -0.0271 0.0497 1.0000
9.250 1.0826 0.03169 0.02211 -0.0250 0.0458 1.0000
9.500 1.0973 0.03300 0.02362 -0.0232 0.0418 1.0000
9.750 1.1101 0.03435 0.02509 -0.0214 0.0384 1.0000
10.000 1.1208 0.03569 0.02651 -0.0195 0.0358 1.0000
10.250 1.1324 0.03771 0.02858 -0.0179 0.0335 1.0000
10.500 1.1469 0.03958 0.03072 -0.0164 0.0318 1.0000
10.750 1.1582 0.04146 0.03285 -0.0147 0.0299 1.0000
11.000 1.1665 0.04329 0.03488 -0.0129 0.0284 1.0000
11.250 1.1719 0.04506 0.03680 -0.0111 0.0270 1.0000
11.500 1.1768 0.04715 0.03905 -0.0094 0.0263 1.0000
11.750 1.1796 0.04947 0.04153 -0.0077 0.0255 1.0000
12.000 1.1791 0.05236 0.04458 -0.0062 0.0248 1.0000
12.250 1.1744 0.05569 0.04815 -0.0047 0.0243 1.0000
12.500 1.1662 0.05906 0.05188 -0.0035 0.0239 1.0000
12.750 1.1552 0.06285 0.05597 -0.0028 0.0236 1.0000
13.000 1.1408 0.06717 0.06058 -0.0027 0.0232 1.0000
13.250 1.1251 0.07190 0.06557 -0.0035 0.0230 1.0000
13.500 1.1067 0.07739 0.07131 -0.0051 0.0228 1.0000
13.750 1.0866 0.08360 0.07775 -0.0078 0.0229 1.0000
14.000 1.0637 0.09089 0.08526 -0.0117 0.0229 1.0000
14.250 1.0398 0.09915 0.09367 -0.0166 0.0232 1.0000
14.500 1.0146 0.10860 0.10328 -0.0225 0.0236 1.0000
14.750 0.9885 0.11929 0.11409 -0.0294 0.0240 1.0000
15.000 0.9647 0.13052 0.12538 -0.0362 0.0244 1.0000
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Polar data table (+)
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