GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 5K AIRFOIL (goe05k-il) Reynolds number: 100,000 Max Cl/Cd: 27.56 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe05k-il-100000-n5.txt Download as CSV file: xf-goe05k-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 5K 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5253 0.10087 0.09587 -0.0112 1.0000 0.0274
-8.000 -0.5252 0.09817 0.09323 -0.0130 1.0000 0.0278
-7.750 -0.5265 0.09550 0.09064 -0.0147 1.0000 0.0280
-7.500 -0.5236 0.09244 0.08761 -0.0178 1.0000 0.0281
-7.250 -0.5175 0.08909 0.08430 -0.0219 1.0000 0.0283
-7.000 -0.5090 0.08559 0.08077 -0.0255 1.0000 0.0284
-6.750 -0.5100 0.08052 0.07580 -0.0210 1.0000 0.0298
-6.500 -0.5035 0.07691 0.07219 -0.0218 1.0000 0.0312
-6.250 -0.4946 0.07321 0.06848 -0.0238 1.0000 0.0328
-6.000 -0.4827 0.06936 0.06458 -0.0266 1.0000 0.0349
-5.750 -0.4615 0.06575 0.06076 -0.0320 1.0000 0.0367
-5.250 -0.4412 0.05766 0.05263 -0.0319 1.0000 0.0398
-5.000 -0.4113 0.05609 0.05048 -0.0352 1.0000 0.0446
-4.750 -0.4041 0.05036 0.04490 -0.0346 1.0000 0.0464
-4.500 -0.3895 0.04683 0.04130 -0.0342 1.0000 0.0498
-4.250 -0.3649 0.04405 0.03802 -0.0352 1.0000 0.0571
-4.000 -0.3500 0.04023 0.03421 -0.0346 1.0000 0.0605
-3.750 -0.3283 0.03742 0.03102 -0.0345 1.0000 0.0706
-3.250 -0.2753 0.03029 0.02313 -0.0319 1.0000 0.0246
-3.000 -0.2472 0.02803 0.02026 -0.0301 1.0000 0.0186
-2.750 -0.2239 0.02536 0.01723 -0.0292 1.0000 0.0171
-2.500 -0.1992 0.02327 0.01470 -0.0281 1.0000 0.0172
-2.250 -0.1738 0.02143 0.01243 -0.0269 1.0000 0.0177
-2.000 -0.1482 0.01945 0.01011 -0.0259 1.0000 0.0175
-1.750 -0.1228 0.01782 0.00823 -0.0248 1.0000 0.0176
-1.500 -0.0984 0.01613 0.00639 -0.0238 1.0000 0.0200
-1.250 -0.0751 0.01509 0.00522 -0.0228 1.0000 0.0238
-1.000 -0.0509 0.01428 0.00419 -0.0218 1.0000 0.0324
-0.750 -0.0008 0.01071 0.00326 -0.0261 1.0000 1.0000
-0.500 0.0219 0.01078 0.00291 -0.0252 1.0000 1.0000
-0.250 0.0445 0.01086 0.00274 -0.0243 1.0000 1.0000
0.000 0.0669 0.01096 0.00265 -0.0234 1.0000 1.0000
0.250 0.0893 0.01107 0.00263 -0.0226 1.0000 1.0000
0.500 0.1116 0.01120 0.00265 -0.0218 1.0000 1.0000
0.750 0.1339 0.01134 0.00274 -0.0210 1.0000 1.0000
1.000 0.1560 0.01150 0.00289 -0.0203 1.0000 1.0000
1.250 0.1781 0.01167 0.00310 -0.0195 1.0000 1.0000
1.500 0.1998 0.01187 0.00336 -0.0187 1.0000 1.0000
1.750 0.2216 0.01209 0.00368 -0.0180 1.0000 1.0000
2.000 0.2612 0.01238 0.00420 -0.0210 0.9925 1.0000
2.250 0.3001 0.01266 0.00486 -0.0238 0.9850 1.0000
2.500 0.3403 0.01289 0.00554 -0.0267 0.9759 1.0000
2.750 0.4660 0.01691 0.00646 -0.0413 0.0185 1.0000
3.000 0.4874 0.01830 0.00799 -0.0395 0.0164 1.0000
3.250 0.5102 0.02014 0.00990 -0.0383 0.0141 1.0000
3.500 0.5363 0.02243 0.01236 -0.0373 0.0139 1.0000
3.750 0.5635 0.02445 0.01472 -0.0361 0.0144 1.0000
4.000 0.5907 0.02697 0.01781 -0.0344 0.0160 1.0000
4.250 0.6145 0.02997 0.02126 -0.0327 0.0177 1.0000
4.500 0.6340 0.03396 0.02563 -0.0312 0.0193 1.0000
4.750 0.6616 0.03581 0.02811 -0.0286 0.0240 1.0000
5.000 0.6808 0.03964 0.03228 -0.0268 0.0290 1.0000
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