GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 590 AIRFOIL (goe590-il) Reynolds number: 200,000 Max Cl/Cd: 67.61 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe590-il-200000.txt Download as CSV file: xf-goe590-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 590 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4155 0.10070 0.09728 -0.0104 1.0000 0.0246
-8.000 -0.4137 0.09840 0.09503 -0.0125 1.0000 0.0248
-7.750 -0.4131 0.09609 0.09277 -0.0146 1.0000 0.0249
-7.500 -0.4069 0.09339 0.09009 -0.0177 1.0000 0.0250
-7.250 -0.3980 0.09045 0.08715 -0.0205 1.0000 0.0251
-7.000 -0.3875 0.08727 0.08397 -0.0228 1.0000 0.0252
-6.750 -0.3759 0.08402 0.08071 -0.0249 1.0000 0.0253
-6.500 -0.3633 0.08066 0.07733 -0.0269 1.0000 0.0253
-6.250 -0.3498 0.07718 0.07382 -0.0286 1.0000 0.0254
-6.000 -0.3447 0.07122 0.06789 -0.0298 1.0000 0.0259
-5.750 -0.3404 0.06649 0.06319 -0.0289 1.0000 0.0266
-5.500 -0.3317 0.06305 0.05977 -0.0283 1.0000 0.0274
-5.250 -0.3204 0.05997 0.05668 -0.0283 1.0000 0.0283
-5.000 -0.3077 0.05700 0.05367 -0.0285 1.0000 0.0294
-4.750 -0.2945 0.05417 0.05080 -0.0286 1.0000 0.0311
-4.500 -0.2822 0.05171 0.04830 -0.0282 1.0000 0.0341
-4.250 -0.2300 0.05008 0.04625 -0.0335 0.9932 0.0378
-4.000 -0.2004 0.04284 0.03893 -0.0383 0.9818 0.0393
-3.750 -0.1664 0.03920 0.03523 -0.0420 0.9687 0.0422
-3.500 -0.1268 0.03598 0.03177 -0.0454 0.9514 0.0472
-3.250 -0.0855 0.03218 0.02755 -0.0484 0.9328 0.0529
-3.000 -0.0521 0.02978 0.02496 -0.0504 0.9106 0.0597
-2.750 -0.0226 0.02757 0.02248 -0.0510 0.8850 0.0720
-2.500 0.0032 0.02549 0.02002 -0.0503 0.8590 0.0815
-2.250 0.0274 0.02391 0.01806 -0.0491 0.8331 0.0937
-2.000 0.0568 0.02034 0.01363 -0.0467 0.8121 0.0654
-1.750 0.0819 0.01830 0.01124 -0.0454 0.7909 0.0612
-1.500 0.1078 0.01633 0.00882 -0.0441 0.7707 0.0584
-1.250 0.1342 0.01506 0.00713 -0.0431 0.7521 0.0592
-1.000 0.1603 0.01476 0.00646 -0.0422 0.7346 0.0641
-0.750 0.1860 0.01355 0.00511 -0.0415 0.7181 0.0663
-0.500 0.2114 0.01301 0.00450 -0.0408 0.7029 0.0700
-0.250 0.2366 0.01265 0.00404 -0.0401 0.6894 0.0743
0.000 0.2614 0.01232 0.00367 -0.0395 0.6771 0.0819
0.250 0.2861 0.01205 0.00332 -0.0387 0.6660 0.0890
0.500 0.3102 0.01164 0.00300 -0.0378 0.6550 0.1202
0.750 0.4079 0.00948 0.00283 -0.0531 0.6423 1.0000
1.000 0.4324 0.00960 0.00284 -0.0524 0.6332 1.0000
1.250 0.4569 0.00974 0.00283 -0.0517 0.6254 1.0000
1.500 0.4816 0.00986 0.00289 -0.0511 0.6166 1.0000
1.750 0.5063 0.01001 0.00297 -0.0505 0.6098 1.0000
2.000 0.5310 0.01015 0.00308 -0.0500 0.6025 1.0000
2.250 0.5558 0.01033 0.00323 -0.0494 0.5962 1.0000
2.500 0.5807 0.01049 0.00339 -0.0488 0.5894 1.0000
2.750 0.6057 0.01070 0.00356 -0.0483 0.5842 1.0000
3.000 0.6306 0.01088 0.00381 -0.0478 0.5779 1.0000
3.250 0.6554 0.01109 0.00397 -0.0472 0.5711 1.0000
3.500 0.6792 0.01120 0.00416 -0.0464 0.5592 1.0000
3.750 0.7030 0.01132 0.00433 -0.0455 0.5472 1.0000
4.000 0.7269 0.01146 0.00452 -0.0447 0.5356 1.0000
4.250 0.7495 0.01152 0.00454 -0.0434 0.5159 1.0000
4.500 0.7704 0.01141 0.00447 -0.0418 0.4821 1.0000
4.750 0.7850 0.01161 0.00419 -0.0389 0.3268 1.0000
5.000 0.7866 0.01479 0.00600 -0.0355 0.0564 1.0000
5.250 0.8046 0.01583 0.00713 -0.0337 0.0433 1.0000
5.500 0.8193 0.01718 0.00857 -0.0316 0.0378 1.0000
5.750 0.8324 0.01878 0.01024 -0.0291 0.0356 1.0000
6.000 0.8499 0.02003 0.01156 -0.0271 0.0339 1.0000
6.250 0.8688 0.02114 0.01272 -0.0256 0.0296 1.0000
6.500 0.8884 0.02279 0.01441 -0.0239 0.0286 1.0000
6.750 0.9111 0.02483 0.01655 -0.0226 0.0284 1.0000
7.000 0.9372 0.02758 0.01960 -0.0212 0.0304 1.0000
7.250 0.9682 0.03119 0.02336 -0.0202 0.0388 1.0000
11.000 0.8679 0.07730 0.07375 0.0061 0.0400 1.0000
11.250 0.8311 0.08538 0.08198 0.0015 0.0416 1.0000
11.500 0.8021 0.09385 0.09054 -0.0034 0.0431 1.0000
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Polar data table (+)
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