GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 562 AIRFOIL (goe562-il) Reynolds number: 100,000 Max Cl/Cd: 40.89 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe562-il-100000.txt Download as CSV file: xf-goe562-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 562 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2195 0.12029 0.11605 -0.0382 1.0000 0.0656
-9.750 -0.2267 0.11886 0.11476 -0.0372 1.0000 0.0668
-9.500 -0.2482 0.11937 0.11549 -0.0339 0.9992 0.0676
-9.250 -0.2369 0.11669 0.11283 -0.0487 0.9873 0.0696
-9.000 -0.2113 0.10964 0.10578 -0.0534 0.9791 0.0709
-8.750 -0.1719 0.10382 0.09990 -0.0542 0.9718 0.0759
-8.500 -0.1510 0.09947 0.09554 -0.0605 0.9552 0.0800
-8.250 -0.1466 0.09638 0.09246 -0.0718 0.9335 0.0833
-8.000 -0.1443 0.09172 0.08776 -0.0838 0.9156 0.0844
-7.750 0.0569 0.06539 0.06129 -0.1000 0.8825 0.1077
-7.500 0.0532 0.06150 0.05730 -0.1111 0.8659 0.1136
-7.250 -0.0522 0.07508 0.07088 -0.1015 0.8856 0.1003
-7.000 -0.0048 0.07056 0.06628 -0.1042 0.8765 0.1095
-6.750 0.0124 0.06580 0.06127 -0.1154 0.8583 0.1161
-6.500 0.0521 0.06202 0.05740 -0.1183 0.8451 0.1244
-6.250 0.0614 0.05868 0.05382 -0.1232 0.8283 0.1325
-6.000 0.0883 0.05597 0.05099 -0.1246 0.8148 0.1414
-5.750 0.0996 0.05336 0.04824 -0.1253 0.8015 0.1517
-5.500 0.1013 0.05173 0.04639 -0.1246 0.7883 0.1642
-5.000 0.1320 0.04769 0.04218 -0.1227 0.7672 0.1898
-4.750 0.1395 0.04600 0.04042 -0.1205 0.7566 0.2038
-4.500 0.1437 0.04493 0.03915 -0.1182 0.7475 0.2287
-4.250 0.1589 0.04295 0.03717 -0.1164 0.7390 0.2474
-4.000 0.1653 0.04198 0.03615 -0.1131 0.7302 0.2785
-3.750 0.1784 0.04042 0.03460 -0.1104 0.7227 0.3122
-3.500 0.1859 0.03915 0.03339 -0.1065 0.7146 0.3462
-3.000 0.2035 0.03655 0.03085 -0.0985 0.7010 0.4275
-2.750 0.2602 0.03141 0.02285 -0.1048 0.6949 0.1294
-2.500 0.2877 0.02938 0.02060 -0.1046 0.6894 0.1238
-2.250 0.3035 0.02924 0.01992 -0.1010 0.6818 0.1146
-2.000 0.3307 0.02757 0.01807 -0.1008 0.6756 0.1113
-1.750 0.3553 0.02672 0.01701 -0.0998 0.6697 0.1088
-1.500 0.3766 0.02617 0.01630 -0.0983 0.6631 0.1073
-1.250 0.4095 0.02551 0.01540 -0.0989 0.6580 0.1092
-1.000 0.4282 0.02535 0.01519 -0.0970 0.6514 0.1110
-0.750 0.4544 0.02499 0.01473 -0.0965 0.6453 0.1127
-0.500 0.4898 0.02453 0.01414 -0.0976 0.6409 0.1146
-0.250 0.5010 0.02450 0.01429 -0.0946 0.6344 0.1174
0.000 0.5249 0.02434 0.01414 -0.0938 0.6289 0.1253
0.250 0.5594 0.02405 0.01375 -0.0949 0.6247 0.1385
0.500 0.5711 0.02441 0.01423 -0.0921 0.6179 0.1529
0.750 0.7910 0.02210 0.01333 -0.1318 0.6082 1.0000
1.000 0.7993 0.02266 0.01391 -0.1282 0.6004 1.0000
1.250 0.8270 0.02273 0.01376 -0.1279 0.5940 1.0000
1.500 0.8403 0.02317 0.01416 -0.1252 0.5863 1.0000
1.750 0.8644 0.02329 0.01412 -0.1244 0.5795 1.0000
2.000 0.8853 0.02362 0.01435 -0.1231 0.5736 1.0000
2.250 0.8992 0.02414 0.01488 -0.1206 0.5672 1.0000
2.500 0.9256 0.02433 0.01494 -0.1203 0.5624 1.0000
2.750 0.9413 0.02488 0.01548 -0.1181 0.5569 1.0000
3.000 0.9564 0.02538 0.01598 -0.1159 0.5506 1.0000
3.250 0.9854 0.02547 0.01594 -0.1160 0.5459 1.0000
3.500 0.9956 0.02617 0.01670 -0.1130 0.5397 1.0000
3.750 1.0145 0.02650 0.01701 -0.1113 0.5336 1.0000
4.000 1.0481 0.02645 0.01681 -0.1122 0.5291 1.0000
4.250 1.0497 0.02737 0.01788 -0.1078 0.5218 1.0000
4.500 1.0740 0.02753 0.01799 -0.1071 0.5164 1.0000
4.750 1.0997 0.02774 0.01814 -0.1067 0.5114 1.0000
5.000 1.1044 0.02855 0.01907 -0.1028 0.5041 1.0000
5.250 1.1362 0.02845 0.01889 -0.1033 0.4991 1.0000
5.500 1.1426 0.02925 0.01979 -0.0997 0.4921 1.0000
5.750 1.1633 0.02947 0.02001 -0.0984 0.4858 1.0000
6.000 1.1948 0.02946 0.01992 -0.0990 0.4807 1.0000
6.250 1.1934 0.03042 0.02104 -0.0941 0.4726 1.0000
6.500 1.2312 0.03011 0.02062 -0.0957 0.4674 1.0000
6.750 1.2270 0.03121 0.02189 -0.0904 0.4593 1.0000
7.000 1.2574 0.03110 0.02174 -0.0907 0.4533 1.0000
7.250 1.2642 0.03187 0.02260 -0.0873 0.4462 1.0000
7.500 1.2818 0.03221 0.02297 -0.0856 0.4395 1.0000
7.750 1.3110 0.03231 0.02303 -0.0859 0.4340 1.0000
8.000 1.3026 0.03358 0.02448 -0.0801 0.4265 1.0000
8.250 1.3401 0.03344 0.02428 -0.0818 0.4217 1.0000
8.500 1.3286 0.03502 0.02603 -0.0758 0.4153 1.0000
8.750 1.3364 0.03594 0.02704 -0.0729 0.4099 1.0000
9.000 1.3771 0.03585 0.02693 -0.0751 0.4061 1.0000
9.250 1.3461 0.03819 0.02945 -0.0664 0.4009 1.0000
9.500 1.3240 0.04014 0.03153 -0.0595 0.3959 1.0000
9.750 1.3477 0.04072 0.03219 -0.0592 0.3927 1.0000
10.000 1.4021 0.04054 0.03202 -0.0634 0.3901 1.0000
10.250 1.1008 0.06394 0.05566 -0.0388 0.3743 1.0000
10.500 1.1498 0.06131 0.05310 -0.0384 0.3747 1.0000
11.500 1.4028 0.04633 0.03835 -0.0430 0.3604 1.0000
11.750 0.7248 0.14072 0.13267 -0.0535 0.4079 1.0000
12.000 0.7419 0.14323 0.13523 -0.0537 0.4067 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 562 AIRFOIL (goe562-il)