GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: GOE 587 AIRFOIL (goe587-il) Reynolds number: 1,000,000 Max Cl/Cd: 77.94 at α=2.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe587-il-1000000.txt Download as CSV file: xf-goe587-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4621 0.08879 0.08721 -0.0132 1.0000 0.0064
-7.750 -0.4649 0.08567 0.08412 -0.0146 1.0000 0.0064
-7.500 -0.4632 0.08217 0.08062 -0.0164 1.0000 0.0064
-7.250 -0.4582 0.07859 0.07704 -0.0184 1.0000 0.0064
-7.000 -0.4608 0.07280 0.07125 -0.0208 1.0000 0.0066
-6.750 -0.4567 0.06868 0.06712 -0.0222 1.0000 0.0068
-6.500 -0.4484 0.06548 0.06389 -0.0230 1.0000 0.0070
-6.250 -0.4389 0.06232 0.06070 -0.0236 1.0000 0.0071
-6.000 -0.4283 0.05925 0.05759 -0.0241 1.0000 0.0073
-5.750 -0.4167 0.05595 0.05424 -0.0244 1.0000 0.0075
-5.500 -0.4044 0.05289 0.05112 -0.0244 1.0000 0.0077
-4.000 -0.2234 0.02390 0.02087 -0.0377 0.9726 0.0095
-3.750 -0.1953 0.01762 0.01404 -0.0371 0.9623 0.0096
-3.500 -0.1609 0.01325 0.00907 -0.0381 0.9527 0.0098
-3.250 -0.1214 0.01138 0.00687 -0.0405 0.9394 0.0104
-3.000 -0.0840 0.00951 0.00464 -0.0424 0.9189 0.0107
-2.750 -0.0559 0.00838 0.00320 -0.0422 0.8940 0.0117
-2.500 -0.0313 0.00777 0.00243 -0.0414 0.8706 0.0144
-2.250 -0.0064 0.00758 0.00212 -0.0406 0.8474 0.0165
-2.000 0.0193 0.00762 0.00209 -0.0401 0.8231 0.0185
-1.750 0.0428 0.00721 0.00155 -0.0391 0.7990 0.0261
-1.500 0.0683 0.00726 0.00149 -0.0385 0.7757 0.0296
-1.250 0.0928 0.00710 0.00124 -0.0378 0.7538 0.0386
-1.000 0.1186 0.00712 0.00118 -0.0374 0.7339 0.0416
-0.750 0.1439 0.00700 0.00097 -0.0368 0.7157 0.0472
-0.500 0.1695 0.00695 0.00086 -0.0363 0.6990 0.0550
-0.250 0.1946 0.00679 0.00079 -0.0358 0.6806 0.1032
0.000 0.2186 0.00663 0.00075 -0.0351 0.6513 0.1855
0.250 0.2358 0.00568 0.00072 -0.0335 0.6246 0.5620
0.500 0.2453 0.00479 0.00079 -0.0293 0.6030 0.9019
0.750 0.2686 0.00487 0.00084 -0.0281 0.5829 0.9295
1.000 0.2925 0.00499 0.00090 -0.0270 0.5628 0.9474
1.250 0.3181 0.00511 0.00096 -0.0264 0.5443 0.9605
1.500 0.3487 0.00528 0.00105 -0.0268 0.5250 0.9724
1.750 0.4000 0.00560 0.00126 -0.0319 0.5021 0.9830
2.000 0.4363 0.00576 0.00131 -0.0340 0.4763 0.9843
2.250 0.4684 0.00601 0.00136 -0.0352 0.4219 0.9856
2.500 0.4953 0.00677 0.00154 -0.0356 0.2831 0.9874
2.750 0.5168 0.00816 0.00202 -0.0352 0.0506 0.9896
3.000 0.5501 0.00837 0.00219 -0.0365 0.0401 0.9920
3.250 0.5806 0.00854 0.00241 -0.0373 0.0379 0.9940
3.500 0.6101 0.00876 0.00265 -0.0378 0.0334 0.9949
3.750 0.6388 0.00920 0.00318 -0.0382 0.0275 0.9959
4.000 0.6687 0.00932 0.00330 -0.0389 0.0265 0.9968
4.250 0.6982 0.00950 0.00349 -0.0395 0.0237 0.9979
4.500 0.7254 0.01014 0.00421 -0.0397 0.0170 0.9991
4.750 0.7557 0.01017 0.00422 -0.0405 0.0155 0.9999
5.000 0.7771 0.01044 0.00452 -0.0393 0.0134 1.0000
5.250 0.7962 0.01075 0.00484 -0.0377 0.0118 1.0000
5.500 0.8032 0.01215 0.00643 -0.0335 0.0098 1.0000
5.750 0.8237 0.01223 0.00652 -0.0322 0.0091 1.0000
6.000 0.8387 0.01287 0.00723 -0.0297 0.0086 1.0000
6.250 0.8550 0.01353 0.00797 -0.0275 0.0079 1.0000
6.500 0.8726 0.01422 0.00873 -0.0256 0.0074 1.0000
6.750 0.8915 0.01488 0.00945 -0.0240 0.0069 1.0000
7.000 0.9107 0.01557 0.01020 -0.0226 0.0063 1.0000
7.250 0.9183 0.02014 0.01516 -0.0188 0.0055 1.0000
15.250 0.8228 0.18271 0.18101 -0.0489 0.0054 1.0000
15.500 0.8279 0.18732 0.18561 -0.0510 0.0054 1.0000
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Polar data table (+)
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