GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 200,000 Max Cl/Cd: 76.23 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe565-il-200000.txt Download as CSV file: xf-goe565-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4301 0.08647 0.08288 -0.0318 1.0000 0.0550
-8.250 -0.4339 0.08351 0.07998 -0.0322 1.0000 0.0570
-8.000 -0.4461 0.08054 0.07709 -0.0327 1.0000 0.0589
-7.750 -0.4604 0.07731 0.07395 -0.0337 1.0000 0.0597
-7.250 -0.4905 0.06604 0.06248 -0.0452 1.0000 0.0617
-7.000 -0.4992 0.05971 0.05611 -0.0453 1.0000 0.0626
-6.750 -0.4864 0.06037 0.05700 -0.0395 1.0000 0.0647
-6.500 -0.4818 0.05824 0.05486 -0.0381 1.0000 0.0667
-6.250 -0.4776 0.05448 0.05100 -0.0389 1.0000 0.0702
-6.000 -0.4766 0.04727 0.04328 -0.0426 1.0000 0.0763
-5.750 -0.4658 0.03894 0.03434 -0.0430 1.0000 0.0602
-5.500 -0.4534 0.03424 0.02934 -0.0426 1.0000 0.0578
-5.250 -0.4369 0.02983 0.02438 -0.0421 1.0000 0.0571
-5.000 -0.4175 0.02630 0.02030 -0.0413 1.0000 0.0573
-4.750 -0.3927 0.02444 0.01791 -0.0410 0.9993 0.0606
-4.500 -0.3594 0.02180 0.01476 -0.0426 0.9966 0.0635
-4.250 -0.3244 0.02089 0.01380 -0.0446 0.9933 0.0681
-4.000 -0.2895 0.01982 0.01245 -0.0460 0.9892 0.0714
-3.750 -0.2515 0.01927 0.01158 -0.0480 0.9852 0.0746
-3.500 -0.2149 0.01777 0.01002 -0.0500 0.9815 0.0778
-3.250 -0.1782 0.01713 0.00933 -0.0518 0.9755 0.0813
-3.000 -0.1372 0.01664 0.00876 -0.0544 0.9714 0.0858
-2.750 -0.1035 0.01606 0.00811 -0.0555 0.9660 0.0885
-2.500 -0.0691 0.01539 0.00753 -0.0570 0.9609 0.0932
-2.250 -0.0292 0.01505 0.00717 -0.0594 0.9573 0.1007
-2.000 0.0016 0.01454 0.00675 -0.0601 0.9510 0.1093
-1.750 0.0372 0.01403 0.00631 -0.0617 0.9458 0.1258
-1.500 0.0738 0.01282 0.00603 -0.0640 0.9425 0.3303
-1.250 0.0999 0.01218 0.00605 -0.0638 0.9351 0.4968
-1.000 0.1331 0.01156 0.00601 -0.0646 0.9300 0.6630
-0.750 0.2160 0.01066 0.00581 -0.0749 0.9322 1.0000
-0.500 0.2672 0.01039 0.00542 -0.0793 0.9282 1.0000
-0.250 0.3002 0.01028 0.00523 -0.0801 0.9194 1.0000
0.000 0.3422 0.01004 0.00492 -0.0826 0.9136 1.0000
0.250 0.3727 0.00990 0.00472 -0.0826 0.9032 1.0000
0.500 0.4082 0.00968 0.00446 -0.0836 0.8951 1.0000
0.750 0.4376 0.00954 0.00427 -0.0834 0.8841 1.0000
1.000 0.4641 0.00946 0.00416 -0.0826 0.8721 1.0000
1.250 0.4912 0.00937 0.00404 -0.0820 0.8598 1.0000
1.500 0.5181 0.00928 0.00393 -0.0812 0.8466 1.0000
1.750 0.5446 0.00922 0.00384 -0.0804 0.8324 1.0000
2.000 0.5705 0.00916 0.00375 -0.0795 0.8158 1.0000
2.250 0.5966 0.00912 0.00367 -0.0785 0.7971 1.0000
2.500 0.6223 0.00912 0.00362 -0.0776 0.7769 1.0000
2.750 0.6475 0.00918 0.00364 -0.0766 0.7556 1.0000
3.000 0.6729 0.00927 0.00367 -0.0756 0.7337 1.0000
3.250 0.6973 0.00940 0.00376 -0.0745 0.7089 1.0000
3.500 0.7209 0.00956 0.00384 -0.0733 0.6791 1.0000
3.750 0.7432 0.00975 0.00392 -0.0717 0.6401 1.0000
4.000 0.7640 0.01004 0.00399 -0.0699 0.5891 1.0000
4.250 0.7836 0.01046 0.00415 -0.0680 0.5348 1.0000
4.500 0.8036 0.01093 0.00440 -0.0663 0.4877 1.0000
4.750 0.8239 0.01141 0.00469 -0.0648 0.4419 1.0000
5.000 0.8435 0.01193 0.00501 -0.0632 0.3925 1.0000
5.250 0.8632 0.01249 0.00539 -0.0617 0.3469 1.0000
5.500 0.8826 0.01310 0.00579 -0.0602 0.3016 1.0000
5.750 0.9026 0.01370 0.00622 -0.0589 0.2595 1.0000
6.000 0.9232 0.01428 0.00666 -0.0576 0.2239 1.0000
6.250 0.9427 0.01497 0.00719 -0.0563 0.1874 1.0000
6.500 0.9604 0.01590 0.00786 -0.0547 0.1311 1.0000
6.750 0.9748 0.01727 0.00878 -0.0526 0.0770 1.0000
7.000 0.9924 0.01828 0.00976 -0.0509 0.0653 1.0000
7.250 1.0091 0.01934 0.01080 -0.0490 0.0587 1.0000
7.500 1.0275 0.02024 0.01180 -0.0474 0.0541 1.0000
7.750 1.0451 0.02121 0.01278 -0.0458 0.0497 1.0000
8.000 1.0594 0.02288 0.01443 -0.0437 0.0467 1.0000
8.250 1.0790 0.02396 0.01564 -0.0423 0.0444 1.0000
8.500 1.0982 0.02513 0.01688 -0.0410 0.0414 1.0000
8.750 1.1174 0.02669 0.01844 -0.0399 0.0389 1.0000
9.000 1.1410 0.02987 0.02172 -0.0395 0.0370 1.0000
9.250 1.1606 0.03133 0.02342 -0.0381 0.0359 1.0000
9.500 1.1785 0.03290 0.02524 -0.0367 0.0344 1.0000
9.750 1.1949 0.03467 0.02723 -0.0351 0.0327 1.0000
10.000 1.2104 0.03696 0.02977 -0.0336 0.0318 1.0000
10.250 1.2236 0.03950 0.03257 -0.0318 0.0312 1.0000
10.500 1.2329 0.04230 0.03566 -0.0297 0.0308 1.0000
10.750 1.2367 0.04549 0.03921 -0.0271 0.0308 1.0000
11.000 1.2389 0.04838 0.04234 -0.0245 0.0303 1.0000
11.250 1.2371 0.05140 0.04558 -0.0219 0.0298 1.0000
11.500 1.2283 0.05444 0.04885 -0.0184 0.0296 1.0000
11.750 1.2150 0.05749 0.05215 -0.0149 0.0296 1.0000
12.000 1.1999 0.06119 0.05605 -0.0124 0.0293 1.0000
12.250 1.1818 0.06445 0.05956 -0.0102 0.0295 1.0000
12.500 1.1005 0.07451 0.07041 -0.0108 0.0333 1.0000
12.750 1.0706 0.08167 0.07778 -0.0140 0.0340 1.0000
13.000 1.0416 0.08991 0.08617 -0.0188 0.0348 1.0000
13.250 1.0146 0.09909 0.09547 -0.0248 0.0357 1.0000
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Polar data table (+)
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