GOE 587 AIRFOIL (goe587-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 587 AIRFOIL (goe587-il) Reynolds number: 100,000 Max Cl/Cd: 50.69 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe587-il-100000-n5.txt Download as CSV file: xf-goe587-il-100000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4508 0.10518 0.10028 -0.0160 1.0000 0.0302
-8.250 -0.4496 0.10223 0.09739 -0.0179 1.0000 0.0303
-8.000 -0.4488 0.09937 0.09459 -0.0196 1.0000 0.0303
-7.750 -0.4460 0.09626 0.09152 -0.0221 1.0000 0.0304
-7.500 -0.4405 0.09301 0.08829 -0.0246 1.0000 0.0304
-7.250 -0.4335 0.08961 0.08489 -0.0267 1.0000 0.0305
-7.000 -0.4307 0.08378 0.07916 -0.0211 1.0000 0.0323
-6.750 -0.4248 0.08054 0.07594 -0.0218 1.0000 0.0336
-6.500 -0.4180 0.07725 0.07266 -0.0233 1.0000 0.0347
-6.250 -0.4097 0.07395 0.06937 -0.0249 1.0000 0.0359
-6.000 -0.3998 0.07063 0.06602 -0.0264 1.0000 0.0368
-5.750 -0.3866 0.06739 0.06271 -0.0285 1.0000 0.0383
-5.250 -0.3502 0.06165 0.05658 -0.0322 1.0000 0.0398
-5.000 -0.3359 0.05824 0.05304 -0.0321 1.0000 0.0398
-4.500 -0.3208 0.04904 0.04390 -0.0294 1.0000 0.0261
-4.000 -0.2791 0.04122 0.03547 -0.0281 1.0000 0.0201
-3.750 -0.2654 0.03793 0.03200 -0.0269 1.0000 0.0197
-3.500 -0.2458 0.03462 0.02842 -0.0262 0.9987 0.0195
-3.000 -0.1750 0.02771 0.02066 -0.0305 0.9828 0.0228
-2.750 -0.1393 0.02453 0.01689 -0.0317 0.9745 0.0234
-2.500 -0.1030 0.02168 0.01345 -0.0328 0.9666 0.0250
-2.250 -0.0671 0.01958 0.01072 -0.0336 0.9574 0.0300
-2.000 -0.0329 0.01786 0.00880 -0.0346 0.9476 0.0349
-1.750 0.0031 0.01662 0.00739 -0.0361 0.9383 0.0476
-1.500 0.0391 0.01556 0.00619 -0.0376 0.9285 0.0666
-1.250 0.0732 0.01485 0.00544 -0.0388 0.9165 0.0844
-1.000 0.1078 0.01427 0.00476 -0.0400 0.9041 0.1011
-0.750 0.1422 0.01373 0.00417 -0.0412 0.8909 0.1337
-0.500 0.2393 0.01135 0.00398 -0.0550 0.8968 1.0000
-0.250 0.2750 0.01130 0.00367 -0.0565 0.8781 1.0000
0.000 0.3081 0.01129 0.00346 -0.0575 0.8584 1.0000
0.250 0.3366 0.01132 0.00333 -0.0575 0.8372 1.0000
0.500 0.3643 0.01137 0.00324 -0.0574 0.8172 1.0000
0.750 0.3904 0.01145 0.00320 -0.0569 0.7967 1.0000
1.000 0.4164 0.01153 0.00318 -0.0564 0.7778 1.0000
1.250 0.4417 0.01164 0.00320 -0.0558 0.7593 1.0000
1.500 0.4662 0.01177 0.00328 -0.0550 0.7410 1.0000
1.750 0.4905 0.01191 0.00337 -0.0542 0.7238 1.0000
2.000 0.5144 0.01207 0.00349 -0.0533 0.7076 1.0000
2.250 0.5381 0.01223 0.00366 -0.0524 0.6922 1.0000
2.500 0.5610 0.01240 0.00387 -0.0513 0.6738 1.0000
2.750 0.5826 0.01257 0.00401 -0.0499 0.6491 1.0000
3.000 0.6035 0.01273 0.00413 -0.0482 0.6196 1.0000
3.250 0.6247 0.01292 0.00430 -0.0467 0.5923 1.0000
3.500 0.6469 0.01312 0.00462 -0.0455 0.5715 1.0000
3.750 0.6679 0.01337 0.00488 -0.0439 0.5429 1.0000
4.000 0.6892 0.01363 0.00522 -0.0424 0.5159 1.0000
4.250 0.7082 0.01397 0.00555 -0.0405 0.4664 1.0000
4.500 0.7158 0.01529 0.00584 -0.0368 0.2579 1.0000
4.750 0.7217 0.01777 0.00722 -0.0338 0.0634 1.0000
5.000 0.7373 0.01908 0.00858 -0.0317 0.0444 1.0000
5.250 0.7516 0.02047 0.01014 -0.0294 0.0348 1.0000
5.500 0.7663 0.02175 0.01159 -0.0271 0.0286 1.0000
5.750 0.7805 0.02329 0.01321 -0.0248 0.0248 1.0000
6.000 0.7989 0.02495 0.01498 -0.0228 0.0223 1.0000
6.250 0.8190 0.02652 0.01663 -0.0215 0.0187 1.0000
6.500 0.8402 0.02936 0.01955 -0.0205 0.0169 1.0000
6.750 0.8640 0.03179 0.02230 -0.0192 0.0164 1.0000
7.000 0.8856 0.03453 0.02542 -0.0178 0.0160 1.0000
7.250 0.9044 0.03739 0.02874 -0.0159 0.0155 1.0000
7.500 0.9205 0.04017 0.03199 -0.0139 0.0147 1.0000
7.750 0.9334 0.04317 0.03546 -0.0117 0.0137 1.0000
8.000 0.9430 0.04652 0.03925 -0.0095 0.0136 1.0000
8.250 0.9491 0.05012 0.04326 -0.0073 0.0137 1.0000
8.500 0.9518 0.05381 0.04733 -0.0050 0.0138 1.0000
8.750 0.9511 0.05752 0.05137 -0.0029 0.0141 1.0000
9.000 0.9468 0.06120 0.05532 -0.0009 0.0143 1.0000
9.250 0.9388 0.06481 0.05917 0.0010 0.0144 1.0000
9.500 0.9264 0.06802 0.06256 0.0031 0.0146 1.0000
9.750 0.9105 0.07166 0.06635 0.0041 0.0147 1.0000
10.000 0.8948 0.07567 0.07050 0.0036 0.0148 1.0000
10.250 0.8787 0.08040 0.07534 0.0016 0.0149 1.0000
10.500 0.8639 0.08574 0.08078 -0.0017 0.0150 1.0000
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Polar data table (+)
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