GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il) Reynolds number: 100,000 Max Cl/Cd: 52.2 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe235-il-100000.txt Download as CSV file: xf-goe235-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 235 (SCHTTE-LANZ) 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4287 0.10099 0.09585 -0.0024 1.0000 0.0643
-7.750 -0.4220 0.09800 0.09290 -0.0040 1.0000 0.0658
-7.500 -0.4163 0.09515 0.09008 -0.0062 1.0000 0.0675
-7.250 -0.4086 0.09247 0.08744 -0.0116 1.0000 0.0695
-7.000 -0.3896 0.09114 0.08598 -0.0297 1.0000 0.0710
-6.750 -0.3849 0.08518 0.08018 -0.0227 1.0000 0.0720
-6.500 -0.3748 0.08160 0.07665 -0.0200 1.0000 0.0737
-6.250 -0.3620 0.07844 0.07352 -0.0211 1.0000 0.0760
-6.000 -0.3469 0.07528 0.07036 -0.0242 1.0000 0.0788
-5.750 -0.3094 0.07416 0.06883 -0.0407 1.0000 0.0837
-5.500 -0.3045 0.06858 0.06348 -0.0375 1.0000 0.0847
-5.250 -0.2962 0.06525 0.06026 -0.0351 1.0000 0.0863
-5.000 -0.2849 0.06259 0.05765 -0.0345 1.0000 0.0888
-4.750 -0.2712 0.06014 0.05519 -0.0354 1.0000 0.0927
-4.500 -0.2450 0.05794 0.05268 -0.0416 1.0000 0.0989
-4.250 -0.2419 0.05496 0.04987 -0.0387 1.0000 0.1006
-4.000 -0.2354 0.05293 0.04789 -0.0367 1.0000 0.1038
-3.750 -0.2106 0.05148 0.04609 -0.0405 1.0000 0.1143
-3.500 -0.2055 0.04867 0.04346 -0.0382 1.0000 0.1167
-3.250 -0.1941 0.04686 0.04169 -0.0373 1.0000 0.1220
-3.000 -0.1731 0.04489 0.03953 -0.0392 1.0000 0.1318
-2.750 -0.1607 0.04311 0.03779 -0.0383 1.0000 0.1377
-2.500 -0.1298 0.04086 0.03544 -0.0414 0.9971 0.1514
-2.250 -0.0822 0.03882 0.03316 -0.0475 0.9904 0.1798
-2.000 -0.0414 0.03611 0.03053 -0.0516 0.9839 0.2033
-0.500 0.2886 0.02297 0.01520 -0.0805 0.9263 0.1426
-0.250 0.3266 0.02185 0.01375 -0.0810 0.9133 0.1276
0.000 0.3612 0.02047 0.01222 -0.0812 0.9001 0.1179
0.250 0.3939 0.01983 0.01137 -0.0806 0.8862 0.1118
0.500 0.4235 0.01886 0.01038 -0.0797 0.8720 0.1094
0.750 0.4510 0.01814 0.00963 -0.0783 0.8571 0.1084
1.000 0.4768 0.01762 0.00908 -0.0766 0.8411 0.1112
1.250 0.5017 0.01717 0.00862 -0.0748 0.8250 0.1132
1.500 0.5256 0.01659 0.00812 -0.0729 0.8094 0.1152
1.750 0.5500 0.01622 0.00778 -0.0713 0.7926 0.1195
2.000 0.5750 0.01598 0.00750 -0.0697 0.7740 0.1283
2.500 0.6233 0.01348 0.00694 -0.0661 0.7325 1.0000
2.750 0.6481 0.01342 0.00664 -0.0642 0.7035 1.0000
3.000 0.6717 0.01334 0.00628 -0.0620 0.6646 1.0000
3.250 0.6944 0.01344 0.00603 -0.0597 0.6100 1.0000
3.500 0.7167 0.01373 0.00595 -0.0579 0.5195 1.0000
3.750 0.7372 0.01454 0.00596 -0.0561 0.4206 1.0000
4.000 0.7605 0.01539 0.00638 -0.0553 0.3740 1.0000
4.250 0.7852 0.01604 0.00680 -0.0547 0.3477 1.0000
4.500 0.8104 0.01658 0.00720 -0.0542 0.3256 1.0000
4.750 0.8357 0.01708 0.00760 -0.0537 0.3079 1.0000
5.000 0.8610 0.01758 0.00802 -0.0532 0.2936 1.0000
5.250 0.8860 0.01806 0.00846 -0.0528 0.2772 1.0000
5.500 0.9105 0.01853 0.00889 -0.0523 0.2583 1.0000
5.750 0.9347 0.01902 0.00934 -0.0517 0.2379 1.0000
6.000 0.9591 0.01944 0.00978 -0.0512 0.2105 1.0000
6.250 0.9837 0.01990 0.01022 -0.0507 0.1750 1.0000
6.500 1.0087 0.02046 0.01067 -0.0500 0.1254 1.0000
6.750 1.0305 0.02151 0.01134 -0.0492 0.0996 1.0000
7.000 1.0514 0.02268 0.01239 -0.0482 0.0921 1.0000
7.250 1.0724 0.02380 0.01355 -0.0472 0.0872 1.0000
7.500 1.0915 0.02517 0.01488 -0.0460 0.0832 1.0000
7.750 1.1112 0.02656 0.01626 -0.0448 0.0809 1.0000
8.000 1.1327 0.02782 0.01761 -0.0437 0.0794 1.0000
8.250 1.1547 0.02920 0.01906 -0.0427 0.0781 1.0000
8.500 1.1774 0.03072 0.02065 -0.0418 0.0769 1.0000
8.750 1.2001 0.03235 0.02237 -0.0409 0.0753 1.0000
9.000 1.2227 0.03418 0.02426 -0.0402 0.0734 1.0000
9.250 1.2458 0.03630 0.02645 -0.0396 0.0722 1.0000
9.500 1.2684 0.03865 0.02897 -0.0389 0.0717 1.0000
9.750 1.2891 0.04118 0.03175 -0.0380 0.0716 1.0000
10.000 1.3074 0.04399 0.03486 -0.0370 0.0714 1.0000
10.250 1.3226 0.04741 0.03860 -0.0359 0.0709 1.0000
10.500 1.3332 0.05142 0.04296 -0.0347 0.0704 1.0000
10.750 1.3405 0.05489 0.04684 -0.0331 0.0705 1.0000
11.000 1.3433 0.05755 0.04998 -0.0309 0.0712 1.0000
11.250 1.3295 0.06135 0.05446 -0.0280 0.0726 1.0000
11.500 1.2899 0.06683 0.06060 -0.0247 0.0743 1.0000
11.750 1.2508 0.07251 0.06665 -0.0234 0.0755 1.0000
12.000 1.2091 0.07984 0.07427 -0.0255 0.0766 1.0000
12.250 1.1603 0.09024 0.08488 -0.0318 0.0779 1.0000
12.500 1.1130 0.10365 0.09842 -0.0408 0.0799 1.0000
12.750 1.1009 0.11128 0.10606 -0.0437 0.0813 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il)