GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 587 AIRFOIL (goe587-il) Reynolds number: 200,000 Max Cl/Cd: 63.46 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe587-il-200000.txt Download as CSV file: xf-goe587-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4540 0.09286 0.08936 -0.0129 1.0000 0.0269
-7.750 -0.4543 0.09015 0.08670 -0.0148 1.0000 0.0276
-7.500 -0.4553 0.08756 0.08416 -0.0176 1.0000 0.0281
-7.250 -0.4495 0.08476 0.08135 -0.0212 1.0000 0.0285
-7.000 -0.4407 0.08179 0.07836 -0.0240 1.0000 0.0288
-6.750 -0.4304 0.07865 0.07518 -0.0262 1.0000 0.0290
-6.500 -0.4193 0.07536 0.07184 -0.0275 1.0000 0.0291
-6.250 -0.4076 0.07195 0.06836 -0.0285 1.0000 0.0292
-6.000 -0.3967 0.06805 0.06442 -0.0291 1.0000 0.0293
-5.750 -0.3994 0.06141 0.05778 -0.0283 1.0000 0.0304
-5.500 -0.3916 0.05808 0.05446 -0.0273 1.0000 0.0313
-5.250 -0.3810 0.05501 0.05135 -0.0267 1.0000 0.0326
-5.000 -0.3678 0.05200 0.04828 -0.0264 1.0000 0.0341
-4.750 -0.3530 0.04895 0.04513 -0.0260 1.0000 0.0364
-4.500 -0.3225 0.04780 0.04355 -0.0261 1.0000 0.0416
-4.250 -0.3163 0.04243 0.03810 -0.0250 1.0000 0.0432
-4.000 -0.3074 0.03970 0.03541 -0.0234 1.0000 0.0451
-3.750 -0.2950 0.03749 0.03311 -0.0217 1.0000 0.0474
-3.500 -0.2809 0.03531 0.03077 -0.0198 1.0000 0.0509
-3.000 -0.2306 0.02970 0.02462 -0.0203 0.9955 0.0592
-2.750 -0.1892 0.02735 0.02171 -0.0228 0.9891 0.0702
-2.500 -0.1054 0.00748 0.00118 -0.0301 0.9708 0.0426
-2.250 -0.1102 0.01972 0.01316 -0.0255 0.9772 0.0485
-2.000 -0.0686 0.01691 0.00961 -0.0265 0.9717 0.0427
-1.750 -0.0310 0.01481 0.00732 -0.0283 0.9644 0.0462
-1.500 0.0104 0.01369 0.00605 -0.0308 0.9574 0.0570
-1.250 0.0485 0.01272 0.00502 -0.0327 0.9477 0.0684
-1.000 0.0874 0.01179 0.00417 -0.0349 0.9382 0.0847
-0.750 0.1266 0.01114 0.00359 -0.0372 0.9270 0.1061
-0.500 0.2272 0.00861 0.00361 -0.0514 0.9431 1.0000
-0.250 0.2755 0.00845 0.00327 -0.0556 0.9245 1.0000
0.000 0.3228 0.00832 0.00297 -0.0596 0.9022 1.0000
0.250 0.3573 0.00830 0.00281 -0.0608 0.8732 1.0000
0.500 0.3832 0.00837 0.00273 -0.0602 0.8437 1.0000
0.750 0.4066 0.00846 0.00268 -0.0590 0.8167 1.0000
1.000 0.4290 0.00856 0.00264 -0.0577 0.7902 1.0000
1.250 0.4501 0.00867 0.00260 -0.0560 0.7605 1.0000
1.500 0.4711 0.00880 0.00258 -0.0543 0.7312 1.0000
1.750 0.4930 0.00894 0.00264 -0.0530 0.7080 1.0000
2.000 0.5153 0.00907 0.00271 -0.0518 0.6885 1.0000
2.250 0.5374 0.00921 0.00279 -0.0505 0.6695 1.0000
2.500 0.5592 0.00932 0.00288 -0.0492 0.6474 1.0000
2.750 0.5806 0.00944 0.00296 -0.0478 0.6234 1.0000
3.000 0.6009 0.00958 0.00303 -0.0461 0.5926 1.0000
3.250 0.6202 0.00978 0.00316 -0.0442 0.5531 1.0000
3.500 0.6390 0.01007 0.00331 -0.0422 0.5072 1.0000
3.750 0.6538 0.01059 0.00343 -0.0395 0.3903 1.0000
4.000 0.6539 0.01309 0.00431 -0.0352 0.0787 1.0000
4.250 0.6724 0.01380 0.00499 -0.0334 0.0601 1.0000
4.500 0.6888 0.01473 0.00602 -0.0312 0.0525 1.0000
4.750 0.7061 0.01552 0.00696 -0.0292 0.0451 1.0000
5.000 0.7177 0.01710 0.00854 -0.0263 0.0408 1.0000
5.250 0.7344 0.01853 0.01002 -0.0240 0.0390 1.0000
5.500 0.7548 0.01978 0.01134 -0.0223 0.0366 1.0000
5.750 0.7758 0.02109 0.01271 -0.0209 0.0324 1.0000
6.000 0.7992 0.02300 0.01475 -0.0196 0.0318 1.0000
6.250 0.8246 0.02554 0.01760 -0.0181 0.0339 1.0000
8.250 0.9404 0.05324 0.04795 -0.0024 0.0391 1.0000
8.500 0.9443 0.05423 0.04945 0.0008 0.0363 1.0000
8.750 0.9444 0.05752 0.05302 0.0029 0.0350 1.0000
9.000 0.9408 0.06098 0.05673 0.0048 0.0339 1.0000
9.250 0.9336 0.06445 0.06039 0.0065 0.0332 1.0000
9.500 0.9211 0.06781 0.06392 0.0084 0.0329 1.0000
9.750 0.9040 0.07101 0.06724 0.0101 0.0328 1.0000
10.000 0.8845 0.07496 0.07130 0.0097 0.0329 1.0000
10.250 0.8646 0.07991 0.07634 0.0071 0.0330 1.0000
10.500 0.8444 0.08640 0.08291 0.0024 0.0335 1.0000
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Polar data table (+)
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