GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 100,000 Max Cl/Cd: 56.36 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe611-il-100000-n5.txt Download as CSV file: xf-goe611-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 611 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1786 0.10888 0.10434 -0.0652 0.9701 0.0373
-9.000 -0.1677 0.10519 0.10066 -0.0712 0.9615 0.0375
-8.750 -0.1576 0.10146 0.09691 -0.0770 0.9520 0.0377
-8.500 -0.1393 0.09567 0.09113 -0.0761 0.9491 0.0388
-8.250 -0.1220 0.09206 0.08751 -0.0764 0.9429 0.0407
-8.000 -0.1085 0.08848 0.08393 -0.0796 0.9353 0.0425
-7.750 -0.1008 0.08520 0.08066 -0.0824 0.9245 0.0440
-7.500 -0.0938 0.08181 0.07726 -0.0860 0.9136 0.0455
-7.250 -0.0822 0.07785 0.07324 -0.0943 0.9026 0.0477
-7.000 -0.0637 0.07365 0.06881 -0.1044 0.8935 0.0487
-6.750 -0.0514 0.07025 0.06521 -0.1078 0.8833 0.0489
-6.500 -0.0391 0.06512 0.06026 -0.1064 0.8789 0.0504
-6.250 -0.0265 0.06230 0.05742 -0.1061 0.8706 0.0520
-6.000 -0.0076 0.05920 0.05422 -0.1080 0.8640 0.0556
-5.750 0.0145 0.05559 0.05011 -0.1137 0.8547 0.0626
-5.500 0.0306 0.05228 0.04694 -0.1131 0.8489 0.0656
-5.250 0.0578 0.05156 0.04553 -0.1159 0.8401 0.0761
-5.000 0.0818 0.04409 0.03789 -0.1164 0.8344 0.0479
-4.750 0.1024 0.04082 0.03447 -0.1169 0.8275 0.0456
-4.500 0.1274 0.03800 0.03127 -0.1173 0.8201 0.0465
-4.250 0.1537 0.03515 0.02805 -0.1177 0.8140 0.0455
-4.000 0.1780 0.03272 0.02525 -0.1175 0.8059 0.0445
-3.750 0.2081 0.03060 0.02262 -0.1178 0.8003 0.0463
-3.500 0.2327 0.02902 0.02064 -0.1171 0.7913 0.0473
-3.250 0.2624 0.02729 0.01848 -0.1172 0.7853 0.0474
-3.000 0.2883 0.02648 0.01725 -0.1164 0.7763 0.0489
-2.750 0.3167 0.02467 0.01524 -0.1167 0.7699 0.0508
-2.500 0.3422 0.02369 0.01407 -0.1161 0.7609 0.0513
-2.250 0.3715 0.02269 0.01285 -0.1161 0.7542 0.0519
-2.000 0.3965 0.02205 0.01214 -0.1155 0.7446 0.0540
-1.750 0.4254 0.02137 0.01131 -0.1154 0.7378 0.0566
-1.500 0.4504 0.02081 0.01068 -0.1148 0.7280 0.0573
-1.250 0.4776 0.02024 0.01003 -0.1144 0.7202 0.0581
-1.000 0.5033 0.01979 0.00954 -0.1139 0.7112 0.0592
-0.750 0.5291 0.01944 0.00910 -0.1134 0.7024 0.0604
-0.500 0.5558 0.01907 0.00866 -0.1131 0.6941 0.0635
-0.250 0.5813 0.01885 0.00839 -0.1126 0.6847 0.0674
0.000 0.6097 0.01862 0.00800 -0.1125 0.6771 0.0704
0.250 0.6350 0.01857 0.00786 -0.1120 0.6670 0.0742
0.500 0.6624 0.01843 0.00764 -0.1118 0.6587 0.0818
0.750 0.6862 0.01714 0.00766 -0.1117 0.6499 0.5335
1.000 0.7191 0.01628 0.00760 -0.1118 0.6411 1.0000
1.250 0.7456 0.01641 0.00751 -0.1115 0.6327 1.0000
1.500 0.7706 0.01660 0.00756 -0.1110 0.6234 1.0000
2.000 0.8220 0.01695 0.00761 -0.1103 0.6054 1.0000
2.250 0.8475 0.01712 0.00765 -0.1099 0.5961 1.0000
2.500 0.8731 0.01730 0.00770 -0.1095 0.5865 1.0000
2.750 0.8972 0.01753 0.00785 -0.1089 0.5761 1.0000
3.000 0.9226 0.01774 0.00794 -0.1085 0.5670 1.0000
3.250 0.9469 0.01798 0.00810 -0.1080 0.5571 1.0000
3.500 0.9712 0.01824 0.00831 -0.1075 0.5480 1.0000
3.750 0.9964 0.01848 0.00847 -0.1071 0.5398 1.0000
4.000 1.0199 0.01878 0.00876 -0.1066 0.5309 1.0000
4.250 1.0452 0.01904 0.00894 -0.1062 0.5231 1.0000
4.500 1.0674 0.01936 0.00928 -0.1054 0.5132 1.0000
4.750 1.0910 0.01965 0.00952 -0.1048 0.5043 1.0000
5.000 1.1133 0.01996 0.00983 -0.1041 0.4944 1.0000
5.250 1.1353 0.02030 0.01018 -0.1032 0.4850 1.0000
5.500 1.1583 0.02061 0.01046 -0.1026 0.4768 1.0000
5.750 1.1793 0.02099 0.01093 -0.1017 0.4680 1.0000
6.000 1.2022 0.02133 0.01124 -0.1010 0.4608 1.0000
6.250 1.2225 0.02174 0.01175 -0.1000 0.4524 1.0000
6.500 1.2445 0.02210 0.01210 -0.0992 0.4452 1.0000
6.750 1.2633 0.02253 0.01266 -0.0980 0.4362 1.0000
7.000 1.2833 0.02293 0.01307 -0.0969 0.4282 1.0000
7.250 1.3011 0.02339 0.01363 -0.0955 0.4191 1.0000
7.500 1.3194 0.02384 0.01416 -0.0942 0.4111 1.0000
7.750 1.3363 0.02433 0.01474 -0.0927 0.4027 1.0000
8.000 1.3522 0.02484 0.01533 -0.0910 0.3943 1.0000
8.250 1.3649 0.02536 0.01590 -0.0888 0.3850 1.0000
8.500 1.3764 0.02599 0.01664 -0.0865 0.3750 1.0000
8.750 1.3888 0.02662 0.01736 -0.0844 0.3662 1.0000
9.000 1.4003 0.02735 0.01819 -0.0823 0.3568 1.0000
9.250 1.4079 0.02821 0.01913 -0.0797 0.3441 1.0000
9.500 1.4120 0.02927 0.02020 -0.0769 0.3279 1.0000
9.750 1.4133 0.03056 0.02146 -0.0740 0.3094 1.0000
10.000 1.4143 0.03205 0.02296 -0.0713 0.2893 1.0000
10.250 1.4159 0.03366 0.02460 -0.0690 0.2711 1.0000
10.500 1.4157 0.03552 0.02645 -0.0668 0.2522 1.0000
10.750 1.4142 0.03761 0.02852 -0.0646 0.2338 1.0000
11.000 1.4111 0.03996 0.03084 -0.0627 0.2143 1.0000
11.250 1.4048 0.04272 0.03354 -0.0608 0.1930 1.0000
11.500 1.3961 0.04586 0.03660 -0.0591 0.1735 1.0000
11.750 1.3876 0.04913 0.03983 -0.0576 0.1583 1.0000
12.000 1.3802 0.05246 0.04315 -0.0565 0.1482 1.0000
12.250 1.3733 0.05588 0.04659 -0.0556 0.1397 1.0000
12.500 1.3660 0.05950 0.05026 -0.0549 0.1325 1.0000
12.750 1.3602 0.06308 0.05391 -0.0545 0.1257 1.0000
13.250 1.3520 0.07016 0.06120 -0.0541 0.1153 1.0000
13.500 1.3472 0.07394 0.06509 -0.0541 0.1095 1.0000
13.750 1.3419 0.07789 0.06914 -0.0543 0.1017 1.0000
14.000 1.3388 0.08165 0.07302 -0.0546 0.0916 1.0000
14.250 1.3259 0.08682 0.07810 -0.0553 0.0535 1.0000
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Polar data table (+)
Polar graphs
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