GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 587 AIRFOIL (goe587-il) Reynolds number: 100,000 Max Cl/Cd: 50.13 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe587-il-100000.txt Download as CSV file: xf-goe587-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 587 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4588 0.09610 0.09123 -0.0129 1.0000 0.0526
-7.750 -0.4615 0.09366 0.08887 -0.0153 1.0000 0.0536
-7.500 -0.4624 0.09153 0.08679 -0.0195 1.0000 0.0545
-7.250 -0.4578 0.08947 0.08469 -0.0241 1.0000 0.0550
-7.000 -0.4502 0.08720 0.08234 -0.0274 1.0000 0.0553
-6.750 -0.4468 0.07951 0.07485 -0.0215 1.0000 0.0581
-6.500 -0.4395 0.07619 0.07155 -0.0213 1.0000 0.0615
-6.250 -0.4301 0.07319 0.06846 -0.0235 1.0000 0.0661
-6.000 -0.4125 0.07276 0.06772 -0.0297 1.0000 0.0687
-5.750 -0.4095 0.06610 0.06124 -0.0269 1.0000 0.0707
-5.500 -0.4005 0.06264 0.05781 -0.0257 1.0000 0.0738
-5.250 -0.3849 0.05986 0.05490 -0.0268 1.0000 0.0800
-5.000 -0.3699 0.05653 0.05138 -0.0280 1.0000 0.0835
-4.750 -0.3603 0.05299 0.04791 -0.0263 1.0000 0.0875
-4.500 -0.3400 0.05093 0.04547 -0.0273 1.0000 0.0966
-4.250 -0.3305 0.04712 0.04180 -0.0255 1.0000 0.1005
-4.000 -0.3133 0.04469 0.03911 -0.0251 1.0000 0.1111
-3.500 -0.2859 0.03941 0.03372 -0.0224 1.0000 0.1296
-3.250 -0.2704 0.03779 0.03177 -0.0212 1.0000 0.1512
-3.000 -0.2601 0.03523 0.02932 -0.0192 1.0000 0.1727
-2.250 -0.2269 0.02919 0.02322 -0.0132 1.0000 0.2774
-2.000 -0.1620 0.02483 0.01678 -0.0131 1.0000 0.0805
-1.750 -0.1411 0.02274 0.01443 -0.0119 1.0000 0.0760
-1.500 -0.1202 0.02180 0.01304 -0.0103 1.0000 0.0813
-1.250 -0.0868 0.01994 0.01091 -0.0114 0.9964 0.0848
-1.000 -0.0393 0.01859 0.00932 -0.0151 0.9871 0.0967
-0.750 0.0064 0.01731 0.00806 -0.0187 0.9772 0.1159
-0.500 0.0518 0.01631 0.00720 -0.0224 0.9664 0.1467
-0.250 0.1325 0.01316 0.00633 -0.0315 0.9738 1.0000
0.000 0.1873 0.01322 0.00606 -0.0373 0.9602 1.0000
0.250 0.2426 0.01319 0.00584 -0.0431 0.9470 1.0000
0.500 0.2990 0.01305 0.00560 -0.0489 0.9338 1.0000
0.750 0.3523 0.01286 0.00536 -0.0540 0.9191 1.0000
1.000 0.4025 0.01266 0.00512 -0.0583 0.9026 1.0000
1.250 0.4396 0.01262 0.00506 -0.0599 0.8808 1.0000
1.500 0.4734 0.01263 0.00504 -0.0607 0.8604 1.0000
1.750 0.4991 0.01277 0.00521 -0.0599 0.8384 1.0000
2.000 0.5241 0.01291 0.00534 -0.0589 0.8188 1.0000
2.250 0.5452 0.01295 0.00532 -0.0566 0.7929 1.0000
2.500 0.5645 0.01300 0.00530 -0.0539 0.7640 1.0000
2.750 0.5857 0.01317 0.00548 -0.0521 0.7423 1.0000
3.000 0.6068 0.01334 0.00566 -0.0503 0.7203 1.0000
3.250 0.6273 0.01346 0.00585 -0.0481 0.6955 1.0000
3.500 0.6479 0.01360 0.00605 -0.0461 0.6717 1.0000
3.750 0.6642 0.01353 0.00594 -0.0428 0.6277 1.0000
4.000 0.6788 0.01354 0.00583 -0.0390 0.5657 1.0000
4.250 0.6896 0.01393 0.00583 -0.0346 0.4468 1.0000
4.500 0.6860 0.01690 0.00673 -0.0296 0.1109 1.0000
4.750 0.7014 0.01821 0.00799 -0.0272 0.0856 1.0000
5.000 0.7161 0.01949 0.00941 -0.0247 0.0766 1.0000
5.250 0.7322 0.02064 0.01063 -0.0223 0.0688 1.0000
5.500 0.7480 0.02226 0.01219 -0.0200 0.0622 1.0000
5.750 0.7705 0.02403 0.01396 -0.0185 0.0596 1.0000
6.000 0.7973 0.02619 0.01619 -0.0174 0.0588 1.0000
6.250 0.8234 0.02851 0.01862 -0.0165 0.0563 1.0000
6.500 0.8477 0.03162 0.02184 -0.0157 0.0537 1.0000
6.750 0.8723 0.03405 0.02505 -0.0133 0.0596 1.0000
7.000 0.8969 0.03737 0.02895 -0.0110 0.0696 1.0000
10.250 0.8432 0.10453 0.09989 -0.0113 0.1559 1.0000
10.500 0.8087 0.11116 0.10631 -0.0193 0.1530 1.0000
10.750 0.6926 0.11342 0.10888 -0.0164 0.1684 1.0000
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Polar data table (+)
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