GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 50,000 Max Cl/Cd: 39.07 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe565-il-50000-n5.txt Download as CSV file: xf-goe565-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 565 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4288 0.09823 0.09094 -0.0316 1.0000 0.0664
-8.750 -0.4312 0.09453 0.08733 -0.0327 1.0000 0.0656
-8.500 -0.4350 0.09107 0.08395 -0.0335 1.0000 0.0661
-8.250 -0.4409 0.08768 0.08067 -0.0342 1.0000 0.0660
-8.000 -0.4489 0.08432 0.07742 -0.0349 1.0000 0.0665
-7.750 -0.4542 0.08035 0.07354 -0.0367 1.0000 0.0666
-7.500 -0.4588 0.07603 0.06929 -0.0389 1.0000 0.0668
-7.250 -0.4623 0.07140 0.06467 -0.0412 1.0000 0.0666
-7.000 -0.4648 0.06649 0.05973 -0.0433 1.0000 0.0665
-6.750 -0.4655 0.06145 0.05458 -0.0452 1.0000 0.0662
-6.500 -0.4634 0.05651 0.04944 -0.0466 1.0000 0.0664
-6.250 -0.4582 0.05171 0.04431 -0.0480 1.0000 0.0684
-6.000 -0.4503 0.04640 0.03832 -0.0495 1.0000 0.0713
-5.750 -0.4371 0.04333 0.03503 -0.0491 1.0000 0.0737
-5.500 -0.4216 0.04060 0.03203 -0.0485 1.0000 0.0757
-5.250 -0.4045 0.03820 0.02923 -0.0481 1.0000 0.0807
-5.000 -0.3849 0.03530 0.02553 -0.0479 1.0000 0.0844
-4.750 -0.3653 0.03312 0.02309 -0.0472 1.0000 0.0865
-4.500 -0.3448 0.03146 0.02121 -0.0464 1.0000 0.0894
-4.250 -0.3235 0.03012 0.01949 -0.0457 1.0000 0.0950
-4.000 -0.3009 0.02869 0.01761 -0.0450 1.0000 0.0986
-3.750 -0.2790 0.02737 0.01616 -0.0442 1.0000 0.1011
-3.500 -0.2568 0.02631 0.01492 -0.0433 1.0000 0.1042
-3.250 -0.2342 0.02542 0.01382 -0.0424 1.0000 0.1081
-3.000 -0.2112 0.02469 0.01290 -0.0417 1.0000 0.1144
-2.750 -0.1892 0.02400 0.01219 -0.0409 1.0000 0.1213
-2.500 -0.1667 0.02345 0.01142 -0.0400 1.0000 0.1273
-2.250 -0.1436 0.02286 0.01082 -0.0394 1.0000 0.1343
-2.000 -0.1205 0.02241 0.01029 -0.0388 1.0000 0.1450
-1.750 -0.0973 0.02192 0.00993 -0.0385 1.0000 0.1670
-1.500 -0.0614 0.02126 0.00977 -0.0408 0.9947 0.2480
-1.250 -0.0257 0.02055 0.00977 -0.0429 0.9878 0.4194
-1.000 0.0025 0.01967 0.00994 -0.0426 0.9809 0.6782
-0.750 0.0508 0.01925 0.00979 -0.0463 0.9730 1.0000
-0.500 0.0889 0.01959 0.00982 -0.0488 0.9646 1.0000
-0.250 0.1278 0.01994 0.00989 -0.0513 0.9562 1.0000
0.000 0.1632 0.02023 0.00999 -0.0532 0.9464 1.0000
0.250 0.1981 0.02051 0.01010 -0.0549 0.9363 1.0000
0.500 0.2333 0.02077 0.01024 -0.0566 0.9258 1.0000
0.750 0.2703 0.02102 0.01038 -0.0586 0.9155 1.0000
1.000 0.3102 0.02121 0.01051 -0.0609 0.9051 1.0000
1.250 0.3444 0.02136 0.01061 -0.0622 0.8920 1.0000
1.750 0.4152 0.02153 0.01077 -0.0647 0.8652 1.0000
2.000 0.4500 0.02158 0.01084 -0.0657 0.8517 1.0000
2.250 0.4833 0.02161 0.01093 -0.0664 0.8379 1.0000
2.500 0.5161 0.02162 0.01100 -0.0668 0.8238 1.0000
2.750 0.5485 0.02159 0.01104 -0.0671 0.8085 1.0000
3.000 0.5810 0.02147 0.01103 -0.0672 0.7914 1.0000
3.250 0.6153 0.02121 0.01085 -0.0673 0.7724 1.0000
3.500 0.6450 0.02096 0.01071 -0.0664 0.7477 1.0000
3.750 0.6757 0.02069 0.01049 -0.0655 0.7207 1.0000
4.000 0.7048 0.02057 0.01042 -0.0646 0.6929 1.0000
4.250 0.7323 0.02059 0.01053 -0.0637 0.6654 1.0000
4.500 0.7580 0.02071 0.01072 -0.0625 0.6353 1.0000
4.750 0.7830 0.02087 0.01093 -0.0613 0.6025 1.0000
5.000 0.8079 0.02109 0.01114 -0.0600 0.5673 1.0000
5.250 0.8327 0.02141 0.01145 -0.0587 0.5320 1.0000
5.500 0.8556 0.02190 0.01188 -0.0572 0.4953 1.0000
5.750 0.8771 0.02250 0.01245 -0.0557 0.4580 1.0000
6.000 0.8972 0.02320 0.01308 -0.0540 0.4197 1.0000
6.250 0.9145 0.02398 0.01374 -0.0520 0.3779 1.0000
6.500 0.9298 0.02485 0.01445 -0.0499 0.3349 1.0000
6.750 0.9446 0.02577 0.01524 -0.0480 0.2935 1.0000
7.000 0.9592 0.02674 0.01610 -0.0461 0.2553 1.0000
7.250 0.9743 0.02777 0.01709 -0.0443 0.2208 1.0000
7.500 0.9877 0.02900 0.01820 -0.0425 0.1880 1.0000
7.750 0.9998 0.03050 0.01961 -0.0406 0.1542 1.0000
8.000 1.0133 0.03213 0.02124 -0.0387 0.1248 1.0000
8.250 1.0276 0.03379 0.02291 -0.0370 0.1052 1.0000
8.500 1.0410 0.03541 0.02451 -0.0352 0.0920 1.0000
8.750 1.0550 0.03719 0.02632 -0.0336 0.0822 1.0000
9.000 1.0689 0.03908 0.02822 -0.0320 0.0754 1.0000
9.250 1.0854 0.04108 0.03047 -0.0306 0.0681 1.0000
9.500 1.0998 0.04314 0.03266 -0.0292 0.0631 1.0000
9.750 1.1162 0.04563 0.03539 -0.0279 0.0587 1.0000
10.000 1.1275 0.04796 0.03804 -0.0263 0.0545 1.0000
10.250 1.1374 0.05035 0.04056 -0.0247 0.0516 1.0000
10.500 1.1486 0.05347 0.04380 -0.0234 0.0495 1.0000
10.750 1.1480 0.05661 0.04742 -0.0210 0.0480 1.0000
11.000 1.1424 0.05995 0.05115 -0.0187 0.0465 1.0000
11.250 1.1341 0.06339 0.05492 -0.0169 0.0453 1.0000
11.500 1.1229 0.06714 0.05896 -0.0155 0.0444 1.0000
11.750 1.1088 0.07131 0.06340 -0.0149 0.0438 1.0000
12.000 1.0909 0.07622 0.06857 -0.0153 0.0437 1.0000
12.250 1.0689 0.08209 0.07469 -0.0171 0.0438 1.0000
12.500 1.0421 0.08944 0.08228 -0.0206 0.0445 1.0000
12.750 1.0126 0.09849 0.09151 -0.0261 0.0456 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 565 AIRFOIL (goe565-il)