GOE 504 AIRFOIL (goe504-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 504 AIRFOIL (goe504-il) Reynolds number: 200,000 Max Cl/Cd: 66.15 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe504-il-200000.txt Download as CSV file: xf-goe504-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 504 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.0314 0.05505 0.05184 -0.1300 0.9270 0.0816
-9.750 -0.3253 0.04185 0.03676 -0.1359 0.9094 0.0491
-9.500 -0.3194 0.03895 0.03370 -0.1339 0.8985 0.0483
-9.250 -0.3040 0.03528 0.02966 -0.1341 0.8938 0.0478
-9.000 -0.3008 0.03359 0.02766 -0.1308 0.8823 0.0481
-8.750 -0.2846 0.03165 0.02534 -0.1297 0.8758 0.0486
-8.500 -0.2717 0.03020 0.02357 -0.1275 0.8667 0.0488
-8.250 -0.2464 0.02773 0.02080 -0.1278 0.8619 0.0491
-8.000 -0.2313 0.02593 0.01884 -0.1258 0.8524 0.0496
-7.750 -0.2010 0.02429 0.01712 -0.1267 0.8475 0.0511
-7.500 -0.1833 0.02350 0.01625 -0.1250 0.8382 0.0525
-7.250 -0.1542 0.02238 0.01499 -0.1252 0.8325 0.0537
-7.000 -0.1341 0.02157 0.01408 -0.1238 0.8240 0.0547
-6.750 -0.1069 0.02067 0.01306 -0.1236 0.8176 0.0559
-6.500 -0.0853 0.02014 0.01242 -0.1224 0.8097 0.0574
-6.250 -0.0627 0.01920 0.01145 -0.1215 0.8027 0.0592
-6.000 -0.0417 0.01860 0.01084 -0.1203 0.7955 0.0610
-5.750 -0.0205 0.01812 0.01030 -0.1191 0.7879 0.0630
-5.500 0.0035 0.01766 0.00975 -0.1183 0.7814 0.0655
-5.250 0.0224 0.01729 0.00933 -0.1166 0.7732 0.0689
-5.000 0.0490 0.01681 0.00876 -0.1163 0.7677 0.0745
-4.750 0.0660 0.01653 0.00847 -0.1142 0.7591 0.0812
-4.500 0.0902 0.01590 0.00785 -0.1136 0.7530 0.1056
-4.250 0.1037 0.01515 0.00771 -0.1113 0.7453 0.2296
-4.000 0.1301 0.01515 0.00765 -0.1109 0.7387 0.2712
-3.750 0.1544 0.01520 0.00765 -0.1101 0.7320 0.2936
-3.500 0.1780 0.01522 0.00761 -0.1092 0.7246 0.3116
-3.250 0.2076 0.01519 0.00745 -0.1095 0.7192 0.3262
-3.000 0.2273 0.01521 0.00746 -0.1079 0.7111 0.3374
-2.750 0.2552 0.01514 0.00732 -0.1078 0.7052 0.3530
-2.500 0.2771 0.01512 0.00731 -0.1067 0.6981 0.3653
-2.250 0.3027 0.01505 0.00719 -0.1062 0.6915 0.3779
-2.000 0.3297 0.01502 0.00708 -0.1061 0.6857 0.3920
-1.750 0.3507 0.01496 0.00707 -0.1048 0.6783 0.4076
-1.500 0.3787 0.01483 0.00693 -0.1048 0.6728 0.4273
-1.250 0.3985 0.01477 0.00697 -0.1033 0.6657 0.4471
-1.000 0.4229 0.01462 0.00689 -0.1026 0.6594 0.4718
-0.750 0.4477 0.01446 0.00681 -0.1020 0.6539 0.5026
-0.500 0.4649 0.01425 0.00685 -0.1000 0.6467 0.5500
-0.250 0.4855 0.01384 0.00682 -0.0983 0.6414 0.6553
0.000 0.5155 0.01359 0.00711 -0.0981 0.6348 0.8200
0.250 0.5888 0.01378 0.00730 -0.1068 0.6277 0.9102
0.500 0.6440 0.01403 0.00742 -0.1121 0.6215 0.9415
0.750 0.6925 0.01426 0.00757 -0.1163 0.6139 0.9610
1.000 0.7471 0.01441 0.00752 -0.1217 0.6081 0.9746
1.250 0.7970 0.01457 0.00766 -0.1265 0.6000 0.9880
1.500 0.8539 0.01461 0.00754 -0.1328 0.5936 0.9980
1.750 0.8764 0.01472 0.00764 -0.1322 0.5868 1.0000
2.000 0.8934 0.01480 0.00765 -0.1304 0.5806 1.0000
2.250 0.9135 0.01491 0.00765 -0.1292 0.5751 1.0000
2.500 0.9258 0.01506 0.00781 -0.1265 0.5682 1.0000
2.750 0.9463 0.01517 0.00782 -0.1253 0.5626 1.0000
3.000 0.9615 0.01535 0.00799 -0.1230 0.5564 1.0000
3.250 0.9776 0.01550 0.00810 -0.1210 0.5500 1.0000
3.500 1.0016 0.01564 0.00812 -0.1204 0.5446 1.0000
3.750 1.0113 0.01585 0.00838 -0.1171 0.5377 1.0000
4.000 1.0311 0.01599 0.00846 -0.1157 0.5319 1.0000
4.250 1.0484 0.01621 0.00864 -0.1139 0.5259 1.0000
4.500 1.0627 0.01639 0.00883 -0.1114 0.5192 1.0000
4.750 1.0875 0.01657 0.00889 -0.1110 0.5138 1.0000
5.000 1.0967 0.01681 0.00920 -0.1076 0.5071 1.0000
5.250 1.1151 0.01700 0.00934 -0.1060 0.5011 1.0000
5.500 1.1346 0.01724 0.00955 -0.1046 0.4953 1.0000
5.750 1.1460 0.01748 0.00983 -0.1016 0.4887 1.0000
6.000 1.1689 0.01767 0.00994 -0.1009 0.4830 1.0000
6.250 1.1784 0.01796 0.01028 -0.0976 0.4765 1.0000
6.500 1.1938 0.01818 0.01050 -0.0955 0.4703 1.0000
6.750 1.2155 0.01846 0.01072 -0.0946 0.4648 1.0000
7.000 1.2217 0.01876 0.01110 -0.0907 0.4584 1.0000
7.250 1.2399 0.01899 0.01128 -0.0891 0.4527 1.0000
7.500 1.2502 0.01932 0.01164 -0.0861 0.4467 1.0000
7.750 1.2588 0.01960 0.01195 -0.0828 0.4404 1.0000
8.000 1.2845 0.01988 0.01214 -0.0828 0.4348 1.0000
8.250 1.2846 0.02030 0.01267 -0.0781 0.4284 1.0000
8.500 1.2992 0.02058 0.01295 -0.0760 0.4221 1.0000
8.750 1.3083 0.02100 0.01339 -0.0731 0.4153 1.0000
9.000 1.3155 0.02137 0.01378 -0.0700 0.4080 1.0000
9.250 1.3239 0.02180 0.01420 -0.0672 0.4003 1.0000
9.500 1.3295 0.02230 0.01473 -0.0640 0.3927 1.0000
9.750 1.3398 0.02281 0.01524 -0.0617 0.3855 1.0000
10.000 1.3457 0.02345 0.01595 -0.0589 0.3785 1.0000
10.250 1.3549 0.02403 0.01651 -0.0566 0.3709 1.0000
10.500 1.3573 0.02486 0.01741 -0.0536 0.3630 1.0000
10.750 1.3677 0.02554 0.01806 -0.0517 0.3561 1.0000
11.000 1.3717 0.02649 0.01913 -0.0492 0.3491 1.0000
11.250 1.3821 0.02724 0.01982 -0.0475 0.3420 1.0000
11.500 1.3822 0.02847 0.02119 -0.0449 0.3343 1.0000
11.750 1.3868 0.02955 0.02223 -0.0428 0.3260 1.0000
12.000 1.3865 0.03099 0.02377 -0.0406 0.3174 1.0000
12.250 1.3895 0.03235 0.02516 -0.0387 0.3093 1.0000
12.500 1.3903 0.03393 0.02680 -0.0369 0.3007 1.0000
12.750 1.3909 0.03566 0.02859 -0.0352 0.2919 1.0000
13.000 1.3889 0.03762 0.03053 -0.0334 0.2820 1.0000
13.250 1.3864 0.03983 0.03282 -0.0320 0.2717 1.0000
13.500 1.3823 0.04227 0.03531 -0.0306 0.2602 1.0000
13.750 1.3760 0.04503 0.03808 -0.0293 0.2475 1.0000
14.000 1.3682 0.04806 0.04111 -0.0282 0.2341 1.0000
14.250 1.3568 0.05159 0.04463 -0.0271 0.2175 1.0000
14.500 1.3484 0.05496 0.04799 -0.0263 0.2035 1.0000
14.750 1.3350 0.05897 0.05196 -0.0256 0.1859 1.0000
15.000 1.3192 0.06336 0.05629 -0.0251 0.1662 1.0000
15.250 1.2959 0.06875 0.06155 -0.0247 0.1405 1.0000
15.500 1.2700 0.07462 0.06729 -0.0246 0.1158 1.0000
15.750 1.2423 0.08088 0.07343 -0.0247 0.0897 1.0000
16.000 1.2140 0.08747 0.07992 -0.0251 0.0600 1.0000
16.250 1.1904 0.09364 0.08604 -0.0257 0.0445 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 504 AIRFOIL (goe504-il)