GOE 375 AIRFOIL (goe375-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 375 AIRFOIL (goe375-il) Reynolds number: 100,000 Max Cl/Cd: 52.05 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe375-il-100000-n5.txt Download as CSV file: xf-goe375-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 375 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3375 0.10869 0.10389 -0.0184 1.0000 0.0216
-8.250 -0.3362 0.10642 0.10170 -0.0197 1.0000 0.0216
-8.000 -0.3333 0.10394 0.09929 -0.0210 1.0000 0.0216
-7.750 -0.3217 0.09803 0.09342 -0.0185 1.0000 0.0219
-7.500 -0.3158 0.09431 0.08974 -0.0177 1.0000 0.0222
-7.250 -0.3126 0.09144 0.08693 -0.0177 1.0000 0.0224
-7.000 -0.3082 0.08854 0.08409 -0.0180 1.0000 0.0228
-6.750 -0.3038 0.08575 0.08137 -0.0187 1.0000 0.0234
-6.500 -0.2986 0.08308 0.07876 -0.0196 1.0000 0.0238
-6.250 -0.2927 0.08041 0.07615 -0.0207 1.0000 0.0245
-6.000 -0.2861 0.07776 0.07355 -0.0219 1.0000 0.0252
-5.750 -0.2790 0.07516 0.07099 -0.0231 1.0000 0.0258
-5.500 -0.2714 0.07264 0.06850 -0.0243 1.0000 0.0263
-5.250 -0.2636 0.07028 0.06616 -0.0252 1.0000 0.0267
-4.750 -0.1959 0.06368 0.05924 -0.0381 0.9893 0.0276
-4.500 -0.1522 0.05997 0.05526 -0.0449 0.9805 0.0278
-4.250 -0.1215 0.05520 0.05043 -0.0488 0.9723 0.0280
-4.000 -0.0972 0.05039 0.04565 -0.0515 0.9652 0.0289
-3.750 -0.0670 0.04688 0.04205 -0.0547 0.9557 0.0306
-3.500 -0.0240 0.04427 0.03910 -0.0588 0.9467 0.0347
-3.250 0.0214 0.04175 0.03610 -0.0626 0.9389 0.0359
-3.000 0.0449 0.03743 0.03186 -0.0645 0.9285 0.0378
-2.750 0.0786 0.03478 0.02900 -0.0665 0.9188 0.0413
-2.500 0.1176 0.03247 0.02626 -0.0686 0.9092 0.0476
-2.250 0.1481 0.03024 0.02377 -0.0699 0.8969 0.0608
-2.000 0.1776 0.02790 0.02130 -0.0712 0.8843 0.0778
-1.500 0.2367 0.02459 0.01753 -0.0731 0.8526 0.1578
-1.250 0.2666 0.02250 0.01525 -0.0731 0.8334 0.1671
-1.000 0.3100 0.02167 0.01351 -0.0710 0.8147 0.0699
-0.750 0.3401 0.02042 0.01187 -0.0701 0.7949 0.0614
-0.500 0.3688 0.02019 0.01124 -0.0690 0.7741 0.0589
-0.250 0.3953 0.01882 0.00975 -0.0682 0.7522 0.0559
0.000 0.4218 0.01801 0.00877 -0.0673 0.7281 0.0543
0.250 0.4481 0.01752 0.00808 -0.0663 0.7011 0.0565
0.500 0.4739 0.01701 0.00738 -0.0653 0.6709 0.0568
0.750 0.4993 0.01651 0.00674 -0.0643 0.6386 0.0560
1.000 0.5241 0.01615 0.00621 -0.0632 0.6061 0.0555
1.250 0.5486 0.01593 0.00580 -0.0622 0.5748 0.0552
1.500 0.5730 0.01584 0.00549 -0.0612 0.5463 0.0553
1.750 0.5980 0.01587 0.00528 -0.0604 0.5202 0.0565
2.000 0.6231 0.01593 0.00511 -0.0597 0.4962 0.0598
2.250 0.6483 0.01603 0.00502 -0.0591 0.4732 0.0610
2.500 0.6734 0.01620 0.00499 -0.0584 0.4514 0.0618
2.750 0.6984 0.01640 0.00502 -0.0578 0.4305 0.0627
3.000 0.7233 0.01662 0.00511 -0.0571 0.4098 0.0642
3.250 0.7479 0.01689 0.00521 -0.0565 0.3899 0.0665
3.500 0.7726 0.01714 0.00538 -0.0559 0.3704 0.0702
3.750 0.8077 0.01582 0.00570 -0.0578 0.3497 1.0000
4.000 0.8317 0.01620 0.00594 -0.0571 0.3336 1.0000
4.250 0.8556 0.01658 0.00619 -0.0564 0.3187 1.0000
4.500 0.8796 0.01698 0.00649 -0.0557 0.3054 1.0000
4.750 0.9034 0.01739 0.00683 -0.0551 0.2933 1.0000
5.000 0.9270 0.01783 0.00718 -0.0544 0.2825 1.0000
5.250 0.9506 0.01828 0.00757 -0.0537 0.2725 1.0000
5.500 0.9744 0.01872 0.00801 -0.0531 0.2635 1.0000
5.750 0.9975 0.01924 0.00848 -0.0524 0.2558 1.0000
6.000 1.0214 0.01971 0.00902 -0.0518 0.2475 1.0000
6.250 1.0441 0.02027 0.00952 -0.0511 0.2405 1.0000
6.500 1.0677 0.02077 0.01016 -0.0505 0.2324 1.0000
6.750 1.0901 0.02138 0.01071 -0.0497 0.2260 1.0000
7.000 1.1133 0.02192 0.01141 -0.0491 0.2186 1.0000
7.250 1.1358 0.02253 0.01205 -0.0484 0.2128 1.0000
7.500 1.1586 0.02319 0.01286 -0.0477 0.2074 1.0000
7.750 1.1813 0.02385 0.01365 -0.0470 0.2019 1.0000
8.000 1.2036 0.02456 0.01442 -0.0463 0.1976 1.0000
8.250 1.2260 0.02533 0.01535 -0.0457 0.1931 1.0000
8.500 1.2479 0.02606 0.01632 -0.0449 0.1881 1.0000
8.750 1.2676 0.02656 0.01688 -0.0439 0.1807 1.0000
9.000 1.2853 0.02682 0.01734 -0.0427 0.1692 1.0000
9.250 1.3038 0.02725 0.01800 -0.0416 0.1588 1.0000
9.500 1.3213 0.02763 0.01854 -0.0404 0.1458 1.0000
9.750 1.3383 0.02815 0.01918 -0.0392 0.1306 1.0000
10.000 1.3533 0.02894 0.02004 -0.0379 0.1123 1.0000
10.500 1.3581 0.03333 0.02390 -0.0332 0.0277 1.0000
10.750 1.3558 0.03562 0.02628 -0.0303 0.0230 1.0000
11.000 1.3526 0.03782 0.02865 -0.0274 0.0207 1.0000
11.250 1.3469 0.04028 0.03133 -0.0249 0.0193 1.0000
11.500 1.3398 0.04302 0.03432 -0.0231 0.0185 1.0000
11.750 1.3292 0.04637 0.03792 -0.0221 0.0179 1.0000
12.000 1.3163 0.05037 0.04216 -0.0220 0.0176 1.0000
12.250 1.3015 0.05511 0.04714 -0.0231 0.0174 1.0000
12.500 1.2838 0.06095 0.05322 -0.0254 0.0173 1.0000
12.750 1.2651 0.06771 0.06019 -0.0288 0.0174 1.0000
13.000 1.2443 0.07542 0.06812 -0.0330 0.0175 1.0000
13.250 1.2239 0.08335 0.07623 -0.0371 0.0178 1.0000
13.500 1.2037 0.09135 0.08441 -0.0410 0.0181 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 375 AIRFOIL (goe375-il)