GOE 484 AIRFOIL (goe484-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 484 AIRFOIL (goe484-il) Reynolds number: 100,000 Max Cl/Cd: 67.3 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe484-il-100000.txt Download as CSV file: xf-goe484-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 484 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3637 0.10331 0.09841 -0.0261 1.0000 0.0758
-7.500 -0.3830 0.10369 0.09890 -0.0239 1.0000 0.0765
-7.250 -0.3969 0.10385 0.09916 -0.0252 1.0000 0.0771
-7.000 -0.4016 0.10344 0.09880 -0.0304 1.0000 0.0774
-6.750 -0.3960 0.09657 0.09197 -0.0217 1.0000 0.0789
-6.500 -0.3924 0.09352 0.08895 -0.0188 1.0000 0.0806
-6.250 -0.3915 0.09119 0.08665 -0.0175 1.0000 0.0825
-6.000 -0.3907 0.08893 0.08441 -0.0172 1.0000 0.0848
-5.750 -0.3881 0.08666 0.08217 -0.0186 1.0000 0.0878
-5.500 -0.3612 0.08504 0.08043 -0.0374 1.0000 0.0916
-5.250 -0.3658 0.08066 0.07615 -0.0303 1.0000 0.0926
-5.000 -0.3648 0.07773 0.07327 -0.0257 1.0000 0.0944
-4.750 -0.3573 0.07522 0.07077 -0.0247 1.0000 0.0978
-4.250 -0.3134 0.06833 0.06375 -0.0366 1.0000 0.1079
-4.000 -0.3054 0.06574 0.06119 -0.0348 1.0000 0.1104
-3.750 -0.2759 0.06298 0.05829 -0.0408 1.0000 0.1188
-3.500 -0.2498 0.05910 0.05432 -0.0457 1.0000 0.1229
-3.250 -0.2021 0.05722 0.05203 -0.0557 1.0000 0.1358
-3.000 -0.1969 0.05342 0.04843 -0.0528 1.0000 0.1375
-2.750 -0.1797 0.05119 0.04625 -0.0523 0.9987 0.1433
-2.500 -0.1084 0.04832 0.04286 -0.0661 0.9928 0.1658
-2.250 -0.0801 0.04488 0.03957 -0.0675 0.9869 0.1702
-2.000 -0.0345 0.04232 0.03685 -0.0734 0.9810 0.1857
-1.750 0.0139 0.04012 0.03439 -0.0798 0.9743 0.2113
-1.500 0.0514 0.03797 0.03224 -0.0829 0.9677 0.2301
-1.250 0.1350 0.03233 0.02493 -0.0932 0.9644 0.1214
-1.000 0.1810 0.02941 0.02157 -0.0967 0.9597 0.1028
-0.750 0.2219 0.02761 0.01935 -0.0992 0.9531 0.0968
-0.500 0.2644 0.02622 0.01746 -0.1017 0.9464 0.0937
-0.250 0.3032 0.02537 0.01645 -0.1038 0.9387 0.0946
0.000 0.3436 0.02487 0.01590 -0.1062 0.9311 0.1008
0.250 0.3785 0.02441 0.01533 -0.1075 0.9220 0.1054
0.500 0.4205 0.02377 0.01483 -0.1101 0.9158 0.1137
0.750 0.4524 0.02342 0.01457 -0.1108 0.9055 0.1289
1.000 0.5009 0.02157 0.01424 -0.1150 0.9019 0.5707
1.250 0.5352 0.02106 0.01412 -0.1160 0.8922 1.0000
1.500 0.5788 0.02114 0.01403 -0.1188 0.8861 1.0000
1.750 0.6076 0.02140 0.01420 -0.1191 0.8749 1.0000
2.000 0.6539 0.02135 0.01408 -0.1223 0.8705 1.0000
2.250 0.6791 0.02166 0.01438 -0.1219 0.8588 1.0000
2.500 0.7113 0.02178 0.01450 -0.1225 0.8489 1.0000
2.750 0.7639 0.02104 0.01380 -0.1260 0.8425 1.0000
3.000 0.8020 0.02052 0.01334 -0.1268 0.8302 1.0000
3.250 0.8504 0.01908 0.01197 -0.1281 0.8158 1.0000
3.500 0.9012 0.01720 0.01013 -0.1291 0.7968 1.0000
3.750 0.9296 0.01666 0.00968 -0.1277 0.7755 1.0000
4.000 0.9624 0.01590 0.00895 -0.1267 0.7508 1.0000
4.250 0.9848 0.01549 0.00857 -0.1241 0.7147 1.0000
4.500 1.0108 0.01511 0.00810 -0.1219 0.6653 1.0000
4.750 1.0304 0.01531 0.00785 -0.1184 0.5771 1.0000
5.000 1.0294 0.01682 0.00808 -0.1120 0.3507 1.0000
5.250 1.0258 0.01984 0.00950 -0.1071 0.1391 1.0000
5.500 1.0388 0.02129 0.01068 -0.1045 0.1165 1.0000
5.750 1.0547 0.02237 0.01179 -0.1022 0.1075 1.0000
6.000 1.0679 0.02372 0.01311 -0.0996 0.1013 1.0000
6.250 1.0854 0.02482 0.01426 -0.0977 0.0952 1.0000
6.500 1.1039 0.02630 0.01564 -0.0960 0.0908 1.0000
6.750 1.1309 0.02822 0.01745 -0.0957 0.0876 1.0000
7.000 1.1583 0.02963 0.01897 -0.0953 0.0835 1.0000
7.250 1.1881 0.03144 0.02084 -0.0954 0.0797 1.0000
7.500 1.2215 0.03389 0.02330 -0.0961 0.0773 1.0000
7.750 1.2519 0.03705 0.02660 -0.0964 0.0743 1.0000
8.000 1.2755 0.03928 0.02922 -0.0952 0.0731 1.0000
8.250 1.2984 0.04224 0.03259 -0.0939 0.0730 1.0000
8.500 1.3166 0.04510 0.03589 -0.0920 0.0723 1.0000
8.750 1.3313 0.04816 0.03940 -0.0897 0.0716 1.0000
9.000 1.3463 0.05262 0.04423 -0.0878 0.0734 1.0000
9.250 1.3510 0.05688 0.04944 -0.0831 0.0825 1.0000
11.750 1.0876 0.13785 0.13328 -0.0871 0.1854 1.0000
12.000 1.0395 0.14522 0.14051 -0.0964 0.1755 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 484 AIRFOIL (goe484-il)