GOE 257 AIRFOIL (goe257-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 257 AIRFOIL (goe257-il) Reynolds number: 200,000 Max Cl/Cd: 74.88 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe257-il-200000.txt Download as CSV file: xf-goe257-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 257 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 -0.3951 0.09878 0.09539 -0.0147 1.0000 0.0299
-8.000 -0.3916 0.09548 0.09213 -0.0166 1.0000 0.0307
-7.750 -0.3903 0.09237 0.08907 -0.0191 1.0000 0.0317
-7.500 -0.3824 0.08890 0.08562 -0.0288 1.0000 0.0328
-7.250 -0.3666 0.08473 0.08139 -0.0381 1.0000 0.0332
-7.000 -0.3581 0.07921 0.07587 -0.0410 1.0000 0.0335
-6.750 -0.3550 0.07518 0.07191 -0.0382 1.0000 0.0342
-6.500 -0.3468 0.07207 0.06883 -0.0375 1.0000 0.0349
-6.250 -0.3369 0.06893 0.06570 -0.0382 1.0000 0.0359
-6.000 -0.3259 0.06571 0.06247 -0.0396 1.0000 0.0371
-5.750 -0.3137 0.06242 0.05913 -0.0413 1.0000 0.0386
-5.500 -0.2985 0.05907 0.05570 -0.0435 1.0000 0.0409
-5.250 -0.2655 0.05345 0.04971 -0.0506 0.9981 0.0439
-5.000 -0.2355 0.04964 0.04597 -0.0542 0.9928 0.0463
-4.750 -0.1932 0.04570 0.04186 -0.0601 0.9864 0.0507
-4.500 -0.1435 0.04053 0.03621 -0.0673 0.9806 0.0570
-4.250 -0.1075 0.03767 0.03332 -0.0711 0.9716 0.0613
-4.000 -0.0663 0.03439 0.02966 -0.0753 0.9598 0.0713
-3.500 0.0007 0.02973 0.02461 -0.0794 0.9277 0.0874
-3.250 0.0296 0.02802 0.02255 -0.0799 0.9078 0.0998
-3.000 0.0556 0.02663 0.02091 -0.0798 0.8872 0.1136
-2.750 0.0796 0.02509 0.01927 -0.0795 0.8683 0.1293
-2.500 0.1043 0.02373 0.01780 -0.0793 0.8504 0.1466
-2.250 0.1295 0.02256 0.01647 -0.0793 0.8336 0.1745
-1.500 0.2295 0.01793 0.01017 -0.0768 0.7951 0.0800
-1.250 0.2587 0.01664 0.00862 -0.0765 0.7832 0.0731
-1.000 0.2877 0.01638 0.00810 -0.0760 0.7717 0.0678
-0.750 0.3157 0.01542 0.00704 -0.0757 0.7612 0.0658
-0.500 0.3435 0.01484 0.00636 -0.0754 0.7509 0.0646
-0.250 0.3718 0.01437 0.00585 -0.0753 0.7399 0.0641
0.000 0.3999 0.01404 0.00548 -0.0752 0.7299 0.0644
0.250 0.4277 0.01383 0.00520 -0.0751 0.7198 0.0657
0.500 0.4558 0.01370 0.00503 -0.0751 0.7083 0.0687
0.750 0.4840 0.01352 0.00480 -0.0751 0.6962 0.0714
1.000 0.5122 0.01347 0.00468 -0.0751 0.6842 0.0754
1.250 0.5403 0.01344 0.00460 -0.0750 0.6728 0.0850
1.500 0.5640 0.01117 0.00449 -0.0741 0.6627 1.0000
1.750 0.5923 0.01130 0.00449 -0.0741 0.6510 1.0000
2.000 0.6203 0.01144 0.00451 -0.0741 0.6392 1.0000
2.250 0.6482 0.01157 0.00454 -0.0741 0.6277 1.0000
2.500 0.6761 0.01172 0.00458 -0.0740 0.6168 1.0000
2.750 0.7040 0.01183 0.00464 -0.0740 0.6051 1.0000
3.000 0.7319 0.01195 0.00472 -0.0740 0.5928 1.0000
3.250 0.7597 0.01207 0.00481 -0.0741 0.5811 1.0000
3.500 0.7875 0.01221 0.00492 -0.0741 0.5704 1.0000
3.750 0.8152 0.01234 0.00500 -0.0741 0.5591 1.0000
4.000 0.8428 0.01246 0.00511 -0.0740 0.5461 1.0000
4.250 0.8703 0.01258 0.00525 -0.0740 0.5326 1.0000
4.500 0.8978 0.01273 0.00540 -0.0740 0.5197 1.0000
4.750 0.9252 0.01290 0.00559 -0.0740 0.5073 1.0000
5.000 0.9524 0.01309 0.00579 -0.0740 0.4944 1.0000
5.250 0.9794 0.01330 0.00602 -0.0739 0.4803 1.0000
5.500 1.0056 0.01353 0.00621 -0.0737 0.4611 1.0000
5.750 1.0315 0.01379 0.00644 -0.0734 0.4379 1.0000
6.000 1.0568 0.01413 0.00671 -0.0731 0.4130 1.0000
6.250 1.0820 0.01445 0.00703 -0.0728 0.3831 1.0000
6.500 1.1075 0.01479 0.00737 -0.0726 0.3542 1.0000
6.750 1.1318 0.01523 0.00773 -0.0723 0.3127 1.0000
7.000 1.1534 0.01605 0.00826 -0.0717 0.2615 1.0000
7.250 1.1744 0.01700 0.00904 -0.0710 0.2229 1.0000
7.500 1.1912 0.01850 0.01006 -0.0699 0.1307 1.0000
7.750 1.1974 0.02143 0.01238 -0.0673 0.0455 1.0000
8.250 1.2326 0.02381 0.01499 -0.0644 0.0334 1.0000
8.500 1.2482 0.02500 0.01632 -0.0628 0.0310 1.0000
8.750 1.2599 0.02644 0.01784 -0.0607 0.0294 1.0000
9.000 1.2641 0.02844 0.01991 -0.0576 0.0281 1.0000
9.250 1.2735 0.02992 0.02148 -0.0551 0.0271 1.0000
9.500 1.2831 0.03131 0.02298 -0.0528 0.0259 1.0000
9.750 1.2893 0.03292 0.02468 -0.0499 0.0251 1.0000
10.000 1.2971 0.03469 0.02653 -0.0473 0.0246 1.0000
10.250 1.3080 0.03658 0.02854 -0.0452 0.0241 1.0000
10.500 1.3226 0.03862 0.03069 -0.0435 0.0238 1.0000
10.750 1.3418 0.04096 0.03318 -0.0424 0.0236 1.0000
11.000 1.3652 0.04397 0.03642 -0.0419 0.0239 1.0000
11.250 1.3801 0.04702 0.03974 -0.0405 0.0239 1.0000
11.500 1.3864 0.04969 0.04264 -0.0383 0.0236 1.0000
11.750 1.3891 0.05226 0.04542 -0.0360 0.0231 1.0000
12.000 1.3864 0.05561 0.04909 -0.0334 0.0233 1.0000
12.250 1.3764 0.06020 0.05411 -0.0307 0.0243 1.0000
12.500 1.3642 0.06468 0.05891 -0.0286 0.0251 1.0000
12.750 1.3502 0.06929 0.06379 -0.0271 0.0258 1.0000
13.000 1.3350 0.07423 0.06896 -0.0264 0.0264 1.0000
13.250 1.3196 0.07991 0.07484 -0.0263 0.0271 1.0000
13.500 1.1623 0.07658 0.07175 -0.0142 0.0251 1.0000
13.750 1.1374 0.08203 0.07740 -0.0147 0.0254 1.0000
14.000 1.1106 0.08743 0.08300 -0.0155 0.0256 1.0000
14.250 1.0810 0.09266 0.08841 -0.0167 0.0258 1.0000
14.500 1.0490 0.09760 0.09352 -0.0181 0.0261 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 257 AIRFOIL (goe257-il)