GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 101 AIRFOIL (goe101-il) Reynolds number: 1,000,000 Max Cl/Cd: 114.73 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe101-il-1000000-n5.txt Download as CSV file: xf-goe101-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 101 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3708 0.08635 0.08482 -0.0192 1.0000 0.0039
-7.750 -0.3704 0.08353 0.08203 -0.0197 1.0000 0.0038
-7.500 -0.3723 0.08052 0.07905 -0.0203 0.9917 0.0038
-7.250 -0.3491 0.07472 0.07324 -0.0294 0.9774 0.0038
-6.750 -0.2839 0.05717 0.05550 -0.0577 0.9271 0.0042
-6.500 -0.2661 0.05201 0.05016 -0.0630 0.8912 0.0044
-6.250 -0.2486 0.04800 0.04596 -0.0662 0.8581 0.0045
-6.000 -0.2298 0.04277 0.04049 -0.0697 0.8267 0.0048
-5.750 -0.2112 0.02995 0.02706 -0.0757 0.8043 0.0057
-5.500 -0.2009 0.01271 0.00803 -0.0766 0.7888 0.0068
-5.250 -0.1761 0.01171 0.00679 -0.0761 0.7644 0.0077
-5.000 -0.1515 0.01048 0.00525 -0.0756 0.7432 0.0088
-4.750 -0.1258 0.00987 0.00442 -0.0752 0.7242 0.0099
-4.500 -0.0985 0.00999 0.00451 -0.0750 0.7056 0.0118
-4.250 -0.0713 0.01008 0.00453 -0.0748 0.6869 0.0133
-4.000 -0.0442 0.01011 0.00447 -0.0745 0.6684 0.0147
-3.500 0.0108 0.01064 0.00490 -0.0742 0.6340 0.0171
-3.250 0.0380 0.01062 0.00476 -0.0741 0.6168 0.0189
-3.000 0.0652 0.01063 0.00466 -0.0739 0.5996 0.0203
-2.750 0.0925 0.01070 0.00464 -0.0737 0.5826 0.0211
-2.500 0.1198 0.01086 0.00471 -0.0735 0.5656 0.0216
-2.250 0.1459 0.01007 0.00375 -0.0733 0.5501 0.0228
-2.000 0.1727 0.00976 0.00337 -0.0731 0.5360 0.0240
-1.750 0.1997 0.00958 0.00312 -0.0729 0.5237 0.0247
-1.500 0.2268 0.00934 0.00281 -0.0727 0.5131 0.0251
-1.250 0.2537 0.00912 0.00251 -0.0725 0.5032 0.0253
-1.000 0.2809 0.00890 0.00224 -0.0723 0.4941 0.0255
-0.750 0.3080 0.00875 0.00204 -0.0722 0.4857 0.0259
-0.500 0.3350 0.00860 0.00182 -0.0720 0.4754 0.0259
-0.250 0.3621 0.00846 0.00160 -0.0718 0.4640 0.0257
0.000 0.3891 0.00837 0.00143 -0.0716 0.4511 0.0255
0.250 0.4163 0.00829 0.00130 -0.0714 0.4414 0.0254
0.500 0.4437 0.00822 0.00119 -0.0713 0.4352 0.0254
0.750 0.4711 0.00818 0.00111 -0.0712 0.4289 0.0255
1.000 0.4985 0.00814 0.00104 -0.0710 0.4219 0.0257
1.250 0.5259 0.00813 0.00100 -0.0709 0.4162 0.0262
1.500 0.5533 0.00812 0.00098 -0.0708 0.4098 0.0272
1.750 0.5805 0.00817 0.00098 -0.0706 0.4004 0.0295
2.000 0.6079 0.00820 0.00100 -0.0705 0.3909 0.0312
2.250 0.6351 0.00825 0.00104 -0.0704 0.3829 0.0329
2.500 0.6623 0.00831 0.00108 -0.0703 0.3730 0.0376
2.750 0.6894 0.00832 0.00116 -0.0702 0.3635 0.0702
3.250 0.7451 0.00667 0.00151 -0.0710 0.3452 1.0000
3.500 0.7719 0.00680 0.00160 -0.0708 0.3338 1.0000
3.750 0.7985 0.00696 0.00171 -0.0706 0.3182 1.0000
4.000 0.8244 0.00720 0.00185 -0.0703 0.2939 1.0000
4.250 0.8473 0.00787 0.00219 -0.0696 0.2249 1.0000
4.500 0.8704 0.00851 0.00255 -0.0690 0.1698 1.0000
4.750 0.8958 0.00883 0.00278 -0.0687 0.1485 1.0000
5.000 0.9190 0.00946 0.00316 -0.0681 0.0964 1.0000
5.250 0.9439 0.00984 0.00345 -0.0677 0.0798 1.0000
5.500 0.9695 0.01011 0.00371 -0.0674 0.0693 1.0000
5.750 0.9911 0.01095 0.00428 -0.0666 0.0197 1.0000
6.000 1.0162 0.01128 0.00461 -0.0662 0.0123 1.0000
6.250 1.0414 0.01158 0.00495 -0.0658 0.0103 1.0000
6.500 1.0659 0.01197 0.00536 -0.0653 0.0083 1.0000
6.750 1.0904 0.01235 0.00578 -0.0649 0.0072 1.0000
7.000 1.1149 0.01270 0.00616 -0.0644 0.0064 1.0000
7.250 1.1389 0.01311 0.00661 -0.0639 0.0056 1.0000
7.500 1.1619 0.01365 0.00719 -0.0632 0.0048 1.0000
7.750 1.1855 0.01408 0.00767 -0.0627 0.0045 1.0000
8.000 1.2086 0.01455 0.00819 -0.0621 0.0041 1.0000
8.250 1.2313 0.01505 0.00875 -0.0614 0.0038 1.0000
8.500 1.2534 0.01560 0.00934 -0.0606 0.0036 1.0000
8.750 1.2745 0.01623 0.01003 -0.0598 0.0033 1.0000
9.000 1.2924 0.01721 0.01113 -0.0584 0.0030 1.0000
9.250 1.3134 0.01778 0.01176 -0.0576 0.0029 1.0000
9.500 1.3335 0.01841 0.01247 -0.0566 0.0027 1.0000
9.750 1.3514 0.01923 0.01338 -0.0554 0.0025 1.0000
10.000 1.3699 0.01993 0.01416 -0.0542 0.0024 1.0000
10.250 1.3861 0.02081 0.01512 -0.0528 0.0023 1.0000
10.500 1.4018 0.02165 0.01604 -0.0513 0.0022 1.0000
10.750 1.4158 0.02257 0.01705 -0.0496 0.0021 1.0000
11.000 1.4281 0.02353 0.01809 -0.0477 0.0020 1.0000
11.250 1.4356 0.02452 0.01917 -0.0450 0.0020 1.0000
11.500 1.4401 0.02563 0.02041 -0.0421 0.0019 1.0000
11.750 1.4432 0.02692 0.02179 -0.0393 0.0019 1.0000
12.000 1.4393 0.02884 0.02385 -0.0363 0.0018 1.0000
12.250 1.4332 0.03121 0.02637 -0.0338 0.0017 1.0000
12.500 1.4345 0.03315 0.02844 -0.0323 0.0017 1.0000
12.750 1.4320 0.03567 0.03110 -0.0311 0.0017 1.0000
13.000 1.4308 0.03829 0.03386 -0.0305 0.0017 1.0000
13.250 1.4266 0.04146 0.03717 -0.0302 0.0017 1.0000
13.500 1.4260 0.04441 0.04026 -0.0303 0.0016 1.0000
13.750 1.4191 0.04829 0.04430 -0.0308 0.0016 1.0000
14.000 1.4153 0.05197 0.04811 -0.0316 0.0016 1.0000
14.250 1.4073 0.05640 0.05270 -0.0328 0.0016 1.0000
14.500 1.4006 0.06082 0.05726 -0.0342 0.0015 1.0000
14.750 1.3889 0.06612 0.06271 -0.0361 0.0015 1.0000
15.000 1.3784 0.07148 0.06822 -0.0382 0.0015 1.0000
15.250 1.3642 0.07763 0.07453 -0.0408 0.0015 1.0000
15.500 1.3514 0.08388 0.08093 -0.0436 0.0014 1.0000
15.750 1.3333 0.09135 0.08856 -0.0471 0.0015 1.0000
16.000 1.3197 0.09832 0.09568 -0.0506 0.0014 1.0000
16.250 1.3024 0.10632 0.10382 -0.0548 0.0015 1.0000
16.500 1.2873 0.11425 0.11189 -0.0591 0.0014 1.0000
16.750 1.2690 0.12319 0.12097 -0.0640 0.0015 1.0000
17.000 1.2500 0.13279 0.13071 -0.0694 0.0015 1.0000
17.250 1.2336 0.14215 0.14020 -0.0748 0.0015 1.0000
17.500 1.2153 0.15232 0.15049 -0.0807 0.0015 1.0000
17.750 1.1971 0.16288 0.16117 -0.0869 0.0015 1.0000
18.000 1.1796 0.17376 0.17216 -0.0933 0.0015 1.0000
18.250 1.1557 0.18745 0.18596 -0.1010 0.0016 1.0000
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Polar data table (+)
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