GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 200,000 Max Cl/Cd: 77.33 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe611-il-200000.txt Download as CSV file: xf-goe611-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1437 0.08894 0.08573 -0.0737 0.9668 0.0341
-8.000 -0.1279 0.08413 0.08091 -0.0849 0.9588 0.0353
-7.750 -0.1110 0.07872 0.07542 -0.0985 0.9468 0.0358
-7.500 -0.0965 0.07318 0.06980 -0.1043 0.9393 0.0361
-7.250 -0.0840 0.06966 0.06635 -0.1018 0.9344 0.0369
-7.000 -0.0650 0.06652 0.06319 -0.1030 0.9299 0.0382
-6.750 -0.0516 0.06339 0.06001 -0.1052 0.9203 0.0398
-6.500 -0.0305 0.05928 0.05577 -0.1102 0.9136 0.0424
-6.250 -0.0055 0.05624 0.05212 -0.1168 0.9025 0.0448
-6.000 0.0048 0.05060 0.04660 -0.1171 0.8956 0.0459
-5.750 0.0212 0.04797 0.04397 -0.1170 0.8883 0.0472
-5.500 0.0398 0.04552 0.04142 -0.1174 0.8808 0.0494
-5.250 0.0711 0.04489 0.04007 -0.1188 0.8727 0.0559
-5.000 0.0845 0.03939 0.03463 -0.1193 0.8653 0.0578
-4.750 0.1046 0.03735 0.03260 -0.1192 0.8579 0.0611
-4.500 0.1302 0.03560 0.03028 -0.1193 0.8498 0.0706
-4.250 0.1511 0.03295 0.02770 -0.1194 0.8425 0.0737
-4.000 0.1768 0.03189 0.02614 -0.1191 0.8342 0.0853
-3.750 0.1991 0.02939 0.02372 -0.1192 0.8268 0.0900
-3.500 0.2230 0.02790 0.02200 -0.1190 0.8183 0.1036
-3.250 0.2478 0.02648 0.02044 -0.1190 0.8108 0.1206
-3.000 0.2713 0.02529 0.01916 -0.1187 0.8016 0.1397
-2.000 0.3931 0.01902 0.01125 -0.1156 0.7688 0.0765
-1.750 0.4232 0.01880 0.01065 -0.1149 0.7607 0.0695
-1.500 0.4504 0.01782 0.00954 -0.1145 0.7513 0.0698
-1.250 0.4776 0.01712 0.00877 -0.1140 0.7421 0.0689
-1.000 0.5055 0.01633 0.00789 -0.1137 0.7338 0.0684
-0.750 0.5317 0.01580 0.00735 -0.1132 0.7236 0.0686
-0.500 0.5590 0.01539 0.00688 -0.1129 0.7148 0.0692
-0.250 0.5852 0.01478 0.00630 -0.1126 0.7052 0.0725
0.000 0.6117 0.01450 0.00600 -0.1123 0.6952 0.0749
0.250 0.6395 0.01429 0.00568 -0.1121 0.6864 0.0780
0.500 0.6663 0.01416 0.00546 -0.1118 0.6756 0.0826
0.750 0.6930 0.01400 0.00528 -0.1115 0.6646 0.0962
1.000 0.7250 0.01181 0.00528 -0.1121 0.6542 1.0000
1.250 0.7512 0.01190 0.00514 -0.1116 0.6435 1.0000
1.500 0.7763 0.01202 0.00513 -0.1110 0.6316 1.0000
1.750 0.8020 0.01216 0.00514 -0.1106 0.6210 1.0000
2.000 0.8285 0.01232 0.00514 -0.1104 0.6120 1.0000
2.250 0.8539 0.01249 0.00523 -0.1100 0.6018 1.0000
2.500 0.8799 0.01267 0.00533 -0.1097 0.5927 1.0000
2.750 0.9062 0.01287 0.00539 -0.1095 0.5839 1.0000
3.000 0.9314 0.01306 0.00556 -0.1091 0.5742 1.0000
3.250 0.9577 0.01328 0.00568 -0.1089 0.5659 1.0000
3.500 0.9829 0.01348 0.00584 -0.1086 0.5564 1.0000
3.750 1.0082 0.01371 0.00601 -0.1082 0.5472 1.0000
4.000 1.0339 0.01394 0.00615 -0.1080 0.5384 1.0000
4.250 1.0584 0.01416 0.00640 -0.1075 0.5292 1.0000
4.500 1.0843 0.01442 0.00658 -0.1073 0.5215 1.0000
4.750 1.1085 0.01465 0.00684 -0.1068 0.5125 1.0000
5.000 1.1336 0.01492 0.00708 -0.1065 0.5045 1.0000
5.250 1.1578 0.01516 0.00733 -0.1060 0.4959 1.0000
5.500 1.1823 0.01544 0.00760 -0.1056 0.4880 1.0000
5.750 1.2064 0.01570 0.00786 -0.1052 0.4799 1.0000
6.000 1.2295 0.01596 0.00817 -0.1045 0.4709 1.0000
6.250 1.2534 0.01624 0.00840 -0.1040 0.4625 1.0000
6.500 1.2752 0.01649 0.00875 -0.1031 0.4531 1.0000
6.750 1.2981 0.01679 0.00905 -0.1025 0.4450 1.0000
7.000 1.3196 0.01707 0.00939 -0.1016 0.4360 1.0000
7.250 1.3407 0.01737 0.00973 -0.1006 0.4266 1.0000
7.500 1.3602 0.01767 0.01003 -0.0993 0.4159 1.0000
7.750 1.3775 0.01798 0.01037 -0.0976 0.4030 1.0000
8.000 1.3932 0.01832 0.01074 -0.0957 0.3890 1.0000
8.250 1.4088 0.01870 0.01116 -0.0938 0.3758 1.0000
8.500 1.4228 0.01913 0.01163 -0.0916 0.3625 1.0000
8.750 1.4326 0.01962 0.01214 -0.0887 0.3477 1.0000
9.000 1.4413 0.02023 0.01274 -0.0858 0.3316 1.0000
9.250 1.4493 0.02096 0.01345 -0.0829 0.3148 1.0000
9.500 1.4567 0.02181 0.01429 -0.0801 0.2948 1.0000
9.750 1.4607 0.02290 0.01532 -0.0771 0.2723 1.0000
10.000 1.4640 0.02417 0.01652 -0.0742 0.2482 1.0000
10.250 1.4652 0.02569 0.01796 -0.0714 0.2244 1.0000
10.500 1.4646 0.02747 0.01963 -0.0687 0.2006 1.0000
10.750 1.4620 0.02954 0.02158 -0.0660 0.1758 1.0000
11.000 1.4595 0.03173 0.02367 -0.0637 0.1584 1.0000
11.250 1.4578 0.03397 0.02585 -0.0617 0.1459 1.0000
11.500 1.4576 0.03617 0.02805 -0.0600 0.1366 1.0000
11.750 1.4592 0.03830 0.03022 -0.0585 0.1294 1.0000
12.000 1.4595 0.04061 0.03258 -0.0571 0.1231 1.0000
12.250 1.4611 0.04288 0.03491 -0.0559 0.1157 1.0000
12.500 1.4597 0.04549 0.03756 -0.0547 0.1086 1.0000
12.750 1.4651 0.04755 0.03971 -0.0539 0.0994 1.0000
13.000 1.4691 0.04981 0.04201 -0.0532 0.0813 1.0000
13.250 1.4455 0.05509 0.04694 -0.0521 0.0424 1.0000
13.500 1.4296 0.05977 0.05159 -0.0513 0.0306 1.0000
13.750 1.4226 0.06358 0.05546 -0.0509 0.0277 1.0000
14.000 1.4181 0.06719 0.05917 -0.0506 0.0261 1.0000
14.250 1.4135 0.07090 0.06302 -0.0505 0.0249 1.0000
14.500 1.4087 0.07474 0.06700 -0.0505 0.0241 1.0000
14.750 1.4045 0.07858 0.07099 -0.0507 0.0235 1.0000
15.000 1.3996 0.08258 0.07515 -0.0510 0.0231 1.0000
15.250 1.3933 0.08685 0.07959 -0.0515 0.0227 1.0000
15.500 1.3859 0.09135 0.08425 -0.0521 0.0223 1.0000
15.750 1.3777 0.09607 0.08913 -0.0530 0.0219 1.0000
16.000 1.3685 0.10100 0.09421 -0.0541 0.0216 1.0000
16.250 1.3584 0.10613 0.09949 -0.0553 0.0212 1.0000
16.500 1.3482 0.11138 0.10488 -0.0567 0.0209 1.0000
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Polar data table (+)
Polar graphs
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