GOE 600 AIRFOIL (goe600-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 600 AIRFOIL (goe600-il) Reynolds number: 50,000 Max Cl/Cd: 32.49 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe600-il-50000.txt Download as CSV file: xf-goe600-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 600 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4309 0.10865 0.10140 -0.0084 1.0000 0.3445
-9.250 -0.6094 0.07752 0.07072 -0.0460 1.0000 0.1521
-9.000 -0.6165 0.07232 0.06547 -0.0466 1.0000 0.1494
-8.750 -0.6285 0.06728 0.06037 -0.0467 1.0000 0.1478
-8.500 -0.6433 0.06240 0.05535 -0.0460 1.0000 0.1463
-8.250 -0.6564 0.05789 0.05061 -0.0445 1.0000 0.1455
-8.000 -0.6653 0.05377 0.04619 -0.0425 1.0000 0.1457
-7.750 -0.6697 0.04996 0.04199 -0.0402 1.0000 0.1466
-7.500 -0.6704 0.04646 0.03801 -0.0378 1.0000 0.1486
-7.250 -0.6670 0.04331 0.03437 -0.0355 1.0000 0.1526
-7.000 -0.6532 0.04115 0.03226 -0.0337 1.0000 0.1597
-6.750 -0.6445 0.03856 0.02900 -0.0316 1.0000 0.1661
-6.500 -0.6290 0.03649 0.02701 -0.0301 1.0000 0.1756
-6.250 -0.6140 0.03431 0.02450 -0.0285 1.0000 0.1854
-6.000 -0.5972 0.03256 0.02254 -0.0270 1.0000 0.1987
-5.750 -0.5788 0.03092 0.02085 -0.0257 1.0000 0.2141
-5.500 -0.5602 0.02959 0.01961 -0.0244 1.0000 0.2347
-5.250 -0.5411 0.02827 0.01835 -0.0230 1.0000 0.2619
-5.000 -0.5226 0.02724 0.01748 -0.0216 1.0000 0.2972
-4.750 -0.5035 0.02638 0.01674 -0.0203 1.0000 0.3368
-4.500 -0.4842 0.02574 0.01628 -0.0190 1.0000 0.3796
-4.250 -0.4663 0.02534 0.01620 -0.0172 1.0000 0.4312
-4.000 -0.4499 0.02515 0.01637 -0.0147 1.0000 0.4945
-3.750 -0.4346 0.02498 0.01651 -0.0115 1.0000 0.5647
-3.500 -0.4202 0.02487 0.01656 -0.0078 1.0000 0.6281
-3.250 -0.4062 0.02483 0.01658 -0.0043 1.0000 0.6814
-3.000 -0.3929 0.02483 0.01664 -0.0005 1.0000 0.7252
-2.750 -0.3793 0.02482 0.01662 0.0031 1.0000 0.7663
-2.500 -0.3662 0.02482 0.01662 0.0067 1.0000 0.8072
-2.250 -0.3514 0.02487 0.01665 0.0101 1.0000 0.8496
-2.000 -0.3255 0.02511 0.01682 0.0114 1.0000 0.8986
-1.750 -0.2382 0.02634 0.01773 0.0012 1.0000 0.9620
-1.500 -0.1326 0.02704 0.01803 -0.0157 1.0000 1.0000
-1.250 -0.1567 0.02599 0.01695 -0.0110 1.0000 1.0000
-1.000 -0.1689 0.02522 0.01607 -0.0076 1.0000 1.0000
-0.750 -0.1596 0.02511 0.01574 -0.0070 1.0000 1.0000
-0.500 -0.1429 0.02532 0.01573 -0.0072 1.0000 1.0000
-0.250 -0.1241 0.02567 0.01589 -0.0074 1.0000 1.0000
0.000 -0.1047 0.02612 0.01616 -0.0077 1.0000 1.0000
0.250 -0.0850 0.02663 0.01652 -0.0079 1.0000 1.0000
0.500 -0.0654 0.02720 0.01697 -0.0082 1.0000 1.0000
0.750 -0.0378 0.02798 0.01762 -0.0099 0.9969 1.0000
1.000 0.0144 0.02931 0.01882 -0.0162 0.9841 1.0000
1.250 0.0626 0.03053 0.01995 -0.0217 0.9706 1.0000
1.500 0.1068 0.03163 0.02098 -0.0262 0.9565 1.0000
1.750 0.1491 0.03268 0.02199 -0.0303 0.9416 1.0000
2.000 0.1911 0.03370 0.02300 -0.0341 0.9260 1.0000
2.250 0.2329 0.03470 0.02400 -0.0377 0.9098 1.0000
2.500 0.2754 0.03565 0.02498 -0.0412 0.8932 1.0000
2.750 0.3074 0.03648 0.02583 -0.0429 0.8749 1.0000
3.000 0.3401 0.03730 0.02670 -0.0445 0.8561 1.0000
3.250 0.3759 0.03807 0.02754 -0.0464 0.8372 1.0000
3.500 0.4157 0.03872 0.02827 -0.0486 0.8185 1.0000
3.750 0.4599 0.03918 0.02884 -0.0511 0.8003 1.0000
4.000 0.4872 0.03983 0.02958 -0.0512 0.7799 1.0000
4.250 0.5198 0.04030 0.03015 -0.0518 0.7597 1.0000
4.500 0.5623 0.04033 0.03034 -0.0532 0.7408 1.0000
4.750 0.6110 0.03990 0.03008 -0.0548 0.7229 1.0000
5.000 0.6388 0.04012 0.03041 -0.0539 0.7022 1.0000
5.250 0.6694 0.04005 0.03050 -0.0531 0.6819 1.0000
5.500 0.7097 0.03924 0.02986 -0.0526 0.6635 1.0000
5.750 0.7522 0.03807 0.02890 -0.0520 0.6456 1.0000
6.000 0.7985 0.03645 0.02745 -0.0514 0.6278 1.0000
6.250 0.8148 0.03688 0.02801 -0.0490 0.6032 1.0000
6.500 0.8662 0.03436 0.02565 -0.0480 0.5821 1.0000
6.750 0.8952 0.03332 0.02469 -0.0457 0.5541 1.0000
7.000 0.9281 0.03198 0.02340 -0.0436 0.5245 1.0000
7.250 0.9585 0.03118 0.02258 -0.0417 0.4941 1.0000
7.500 0.9873 0.03090 0.02220 -0.0400 0.4624 1.0000
7.750 1.0147 0.03123 0.02235 -0.0385 0.4298 1.0000
8.000 1.0354 0.03227 0.02332 -0.0367 0.3964 1.0000
8.250 1.0531 0.03350 0.02449 -0.0346 0.3625 1.0000
8.500 1.0685 0.03455 0.02542 -0.0321 0.3280 1.0000
8.750 1.0810 0.03525 0.02589 -0.0294 0.2936 1.0000
9.000 1.0924 0.03634 0.02686 -0.0268 0.2628 1.0000
9.250 1.1030 0.03781 0.02823 -0.0243 0.2337 1.0000
9.500 1.1130 0.03965 0.02997 -0.0217 0.2055 1.0000
9.750 1.1240 0.04191 0.03215 -0.0194 0.1806 1.0000
10.000 1.1383 0.04428 0.03446 -0.0177 0.1612 1.0000
10.250 1.1513 0.04694 0.03723 -0.0159 0.1475 1.0000
10.500 1.1618 0.04999 0.04053 -0.0141 0.1382 1.0000
10.750 1.1802 0.05277 0.04323 -0.0133 0.1289 1.0000
11.000 1.1742 0.05625 0.04728 -0.0100 0.1252 1.0000
11.250 1.1719 0.05971 0.05108 -0.0075 0.1216 1.0000
11.500 1.1904 0.06300 0.05427 -0.0071 0.1157 1.0000
11.750 1.1708 0.06673 0.05843 -0.0036 0.1149 1.0000
12.000 1.1475 0.07054 0.06256 -0.0003 0.1145 1.0000
12.250 1.1224 0.07493 0.06722 0.0019 0.1145 1.0000
12.500 1.0955 0.07997 0.07249 0.0028 0.1148 1.0000
12.750 1.0681 0.08573 0.07844 0.0026 0.1153 1.0000
13.000 1.0413 0.09222 0.08507 0.0012 0.1160 1.0000
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Polar data table (+)
Polar graphs
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