GOE 257 AIRFOIL (goe257-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: GOE 257 AIRFOIL (goe257-il) Reynolds number: 500,000 Max Cl/Cd: 95.99 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe257-il-500000.txt Download as CSV file: xf-goe257-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 257 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4022 0.09628 0.09414 -0.0131 1.0000 0.0176
-8.000 -0.3981 0.09289 0.09077 -0.0149 1.0000 0.0180
-7.750 -0.3957 0.08951 0.08742 -0.0168 1.0000 0.0186
-7.500 -0.3891 0.08547 0.08340 -0.0223 1.0000 0.0195
-7.250 -0.3750 0.08033 0.07825 -0.0314 1.0000 0.0198
-7.000 -0.3596 0.07530 0.07319 -0.0376 1.0000 0.0199
-6.750 -0.3449 0.07063 0.06847 -0.0418 1.0000 0.0199
-6.500 -0.3317 0.06624 0.06403 -0.0446 0.9999 0.0199
-6.250 -0.3138 0.05999 0.05776 -0.0488 0.9953 0.0204
-6.000 -0.2889 0.05702 0.05476 -0.0518 0.9877 0.0210
-5.750 -0.2590 0.05323 0.05088 -0.0565 0.9775 0.0219
-5.500 -0.2274 0.04894 0.04644 -0.0613 0.9631 0.0234
-5.250 -0.1841 0.04408 0.04115 -0.0659 0.9449 0.0256
-5.000 -0.1592 0.04039 0.03714 -0.0667 0.9229 0.0257
-4.750 -0.1439 0.03483 0.03142 -0.0677 0.8983 0.0266
-4.500 -0.1232 0.03303 0.02946 -0.0676 0.8670 0.0273
-4.250 -0.0996 0.03112 0.02728 -0.0677 0.8293 0.0283
-4.000 -0.0733 0.02906 0.02487 -0.0678 0.7953 0.0301
-3.750 -0.0402 0.02865 0.02392 -0.0672 0.7718 0.0331
-3.500 -0.0141 0.02375 0.01862 -0.0686 0.7570 0.0348
-3.250 0.0128 0.02243 0.01716 -0.0691 0.7443 0.0360
-3.000 0.0410 0.02127 0.01578 -0.0695 0.7340 0.0381
-2.750 0.0711 0.02107 0.01530 -0.0694 0.7246 0.0423
-2.500 0.1001 0.01858 0.01248 -0.0702 0.7167 0.0456
-2.250 0.1281 0.01779 0.01161 -0.0706 0.7087 0.0484
-2.000 0.1577 0.01899 0.01263 -0.0703 0.7008 0.0550
-1.750 0.1858 0.01626 0.00978 -0.0713 0.6941 0.0609
-1.500 0.2146 0.01566 0.00905 -0.0717 0.6874 0.0722
-1.250 0.2432 0.01520 0.00848 -0.0721 0.6803 0.0856
-1.000 0.2715 0.01456 0.00782 -0.0726 0.6729 0.1009
-0.500 0.3338 0.01206 0.00483 -0.0713 0.6573 0.0463
-0.250 0.3628 0.01201 0.00469 -0.0713 0.6491 0.0437
0.000 0.3916 0.01149 0.00418 -0.0715 0.6409 0.0424
0.250 0.4203 0.01111 0.00374 -0.0716 0.6327 0.0420
0.500 0.4493 0.01081 0.00344 -0.0718 0.6232 0.0422
0.750 0.4782 0.01062 0.00322 -0.0720 0.6145 0.0425
1.000 0.5070 0.01048 0.00304 -0.0722 0.6053 0.0429
1.250 0.5360 0.01033 0.00288 -0.0725 0.5952 0.0441
1.500 0.5648 0.01025 0.00275 -0.0727 0.5842 0.0460
1.750 0.5935 0.01023 0.00267 -0.0729 0.5723 0.0473
2.000 0.6221 0.01022 0.00263 -0.0730 0.5592 0.0502
2.250 0.6501 0.00906 0.00277 -0.0739 0.5459 0.6136
2.500 0.6743 0.00810 0.00265 -0.0729 0.5305 1.0000
2.750 0.7026 0.00823 0.00267 -0.0730 0.5113 1.0000
3.000 0.7308 0.00839 0.00273 -0.0732 0.4949 1.0000
3.250 0.7589 0.00855 0.00282 -0.0734 0.4816 1.0000
3.500 0.7870 0.00873 0.00294 -0.0735 0.4691 1.0000
3.750 0.8150 0.00892 0.00307 -0.0737 0.4565 1.0000
4.000 0.8428 0.00913 0.00321 -0.0738 0.4442 1.0000
4.250 0.8708 0.00930 0.00337 -0.0740 0.4324 1.0000
4.500 0.8987 0.00950 0.00353 -0.0742 0.4193 1.0000
4.750 0.9261 0.00975 0.00371 -0.0743 0.4008 1.0000
5.000 0.9537 0.00997 0.00388 -0.0744 0.3795 1.0000
5.250 0.9810 0.01022 0.00408 -0.0745 0.3571 1.0000
5.500 1.0075 0.01062 0.00431 -0.0746 0.3118 1.0000
5.750 1.0315 0.01146 0.00477 -0.0745 0.2452 1.0000
6.000 1.0570 0.01202 0.00518 -0.0744 0.2183 1.0000
6.250 1.0829 0.01248 0.00558 -0.0744 0.1982 1.0000
6.500 1.1084 0.01298 0.00597 -0.0743 0.1721 1.0000
6.750 1.1262 0.01470 0.00702 -0.0735 0.0593 1.0000
7.250 1.1727 0.01626 0.00841 -0.0725 0.0208 1.0000
7.500 1.1967 0.01686 0.00911 -0.0720 0.0192 1.0000
7.750 1.2194 0.01759 0.00992 -0.0714 0.0176 1.0000
8.000 1.2397 0.01859 0.01101 -0.0704 0.0160 1.0000
8.250 1.2584 0.01970 0.01226 -0.0692 0.0152 1.0000
8.500 1.2781 0.02058 0.01326 -0.0681 0.0147 1.0000
8.750 1.2958 0.02160 0.01437 -0.0667 0.0143 1.0000
9.000 1.3115 0.02272 0.01560 -0.0651 0.0138 1.0000
9.250 1.3251 0.02395 0.01692 -0.0632 0.0134 1.0000
9.500 1.3365 0.02525 0.01832 -0.0609 0.0130 1.0000
9.750 1.3463 0.02657 0.01974 -0.0585 0.0126 1.0000
10.000 1.3538 0.02796 0.02120 -0.0558 0.0122 1.0000
10.250 1.3554 0.02956 0.02286 -0.0524 0.0118 1.0000
10.500 1.3551 0.03174 0.02509 -0.0489 0.0114 1.0000
10.750 1.3601 0.03417 0.02761 -0.0463 0.0112 1.0000
11.000 1.3733 0.03695 0.03050 -0.0447 0.0111 1.0000
11.250 1.3812 0.03889 0.03259 -0.0427 0.0109 1.0000
11.500 1.3865 0.04066 0.03453 -0.0406 0.0108 1.0000
11.750 1.3913 0.04271 0.03677 -0.0387 0.0107 1.0000
12.000 1.3955 0.04515 0.03938 -0.0370 0.0107 1.0000
12.250 1.3975 0.04792 0.04235 -0.0352 0.0107 1.0000
12.500 1.3970 0.05115 0.04579 -0.0336 0.0107 1.0000
12.750 1.4128 0.05461 0.04934 -0.0324 0.0112 1.0000
13.000 1.4079 0.05698 0.05190 -0.0311 0.0114 1.0000
13.250 1.2560 0.05645 0.05173 -0.0182 0.0112 1.0000
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Polar data table (+)
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