GOE 326 (PFALZ 55) AIRFOIL (goe326-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 326 (PFALZ 55) AIRFOIL (goe326-il) Reynolds number: 100,000 Max Cl/Cd: 59.06 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe326-il-100000-n5.txt Download as CSV file: xf-goe326-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 326 (PFALZ 55) 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
------ -------- --------- --------- -------- -------- --------
-7.250 -0.3195 0.09211 0.08770 -0.0381 1.0000 0.0433
-7.000 -0.3212 0.08718 0.08286 -0.0333 1.0000 0.0442
-6.750 -0.3217 0.08451 0.08026 -0.0296 1.0000 0.0452
-6.500 -0.3162 0.08168 0.07747 -0.0299 0.9974 0.0461
-6.250 -0.2889 0.07718 0.07292 -0.0360 0.9888 0.0480
-6.000 -0.2580 0.07252 0.06819 -0.0438 0.9792 0.0510
-5.750 -0.1948 0.06716 0.06234 -0.0650 0.9663 0.0561
-5.500 -0.1824 0.06167 0.05702 -0.0641 0.9592 0.0577
-5.250 -0.1550 0.05783 0.05312 -0.0672 0.9515 0.0598
-5.000 -0.1072 0.05436 0.04916 -0.0770 0.9406 0.0695
-4.750 -0.0849 0.04957 0.04440 -0.0791 0.9334 0.0721
-4.500 -0.0584 0.04654 0.04129 -0.0810 0.9247 0.0751
-4.000 0.0222 0.03688 0.03039 -0.0895 0.9085 0.0557
-3.750 0.0491 0.03377 0.02711 -0.0907 0.8993 0.0545
-3.500 0.0807 0.03100 0.02399 -0.0921 0.8917 0.0540
-3.250 0.1113 0.02923 0.02165 -0.0926 0.8814 0.0558
-2.500 0.1976 0.02428 0.01564 -0.0932 0.8523 0.0561
-2.250 0.2254 0.02287 0.01396 -0.0930 0.8419 0.0563
-2.000 0.2535 0.02154 0.01238 -0.0929 0.8319 0.0568
-1.750 0.2807 0.02047 0.01115 -0.0927 0.8208 0.0583
-1.500 0.3074 0.01979 0.01039 -0.0923 0.8087 0.0614
-1.250 0.3347 0.01910 0.00955 -0.0918 0.7967 0.0631
-1.000 0.3621 0.01841 0.00870 -0.0913 0.7848 0.0637
-0.750 0.3893 0.01780 0.00796 -0.0907 0.7728 0.0645
-0.500 0.4162 0.01727 0.00734 -0.0900 0.7604 0.0657
-0.250 0.4426 0.01684 0.00683 -0.0894 0.7470 0.0673
0.000 0.4690 0.01649 0.00640 -0.0888 0.7335 0.0692
0.250 0.4953 0.01613 0.00599 -0.0882 0.7197 0.0719
0.500 0.5216 0.01591 0.00572 -0.0876 0.7056 0.0786
0.750 0.5484 0.01574 0.00548 -0.0871 0.6911 0.0876
1.000 0.5754 0.01561 0.00528 -0.0867 0.6759 0.0987
1.250 0.6023 0.01525 0.00519 -0.0864 0.6604 0.1779
1.750 0.6565 0.01367 0.00506 -0.0856 0.6283 1.0000
2.000 0.6827 0.01382 0.00505 -0.0851 0.6122 1.0000
2.500 0.7349 0.01418 0.00510 -0.0841 0.5811 1.0000
3.000 0.7869 0.01463 0.00531 -0.0831 0.5531 1.0000
3.250 0.8128 0.01490 0.00548 -0.0827 0.5400 1.0000
3.500 0.8387 0.01519 0.00570 -0.0823 0.5276 1.0000
3.750 0.8644 0.01550 0.00594 -0.0819 0.5153 1.0000
4.000 0.8900 0.01582 0.00619 -0.0814 0.5030 1.0000
4.250 0.9154 0.01616 0.00648 -0.0810 0.4907 1.0000
4.500 0.9407 0.01649 0.00679 -0.0805 0.4778 1.0000
4.750 0.9658 0.01681 0.00713 -0.0800 0.4646 1.0000
5.000 0.9909 0.01714 0.00749 -0.0796 0.4520 1.0000
5.250 1.0158 0.01748 0.00785 -0.0791 0.4398 1.0000
5.500 1.0405 0.01782 0.00821 -0.0786 0.4280 1.0000
5.750 1.0650 0.01817 0.00857 -0.0780 0.4162 1.0000
6.000 1.0894 0.01852 0.00903 -0.0775 0.4038 1.0000
6.250 1.1137 0.01888 0.00947 -0.0770 0.3921 1.0000
6.500 1.1374 0.01926 0.00992 -0.0764 0.3794 1.0000
6.750 1.1606 0.01965 0.01039 -0.0757 0.3654 1.0000
7.000 1.1835 0.02007 0.01088 -0.0749 0.3519 1.0000
7.250 1.2063 0.02051 0.01141 -0.0742 0.3392 1.0000
7.500 1.2280 0.02098 0.01197 -0.0733 0.3232 1.0000
7.750 1.2471 0.02154 0.01250 -0.0721 0.2982 1.0000
8.000 1.2650 0.02219 0.01315 -0.0708 0.2669 1.0000
8.250 1.2827 0.02294 0.01387 -0.0696 0.2367 1.0000
8.500 1.2988 0.02383 0.01470 -0.0682 0.2030 1.0000
8.750 1.3130 0.02491 0.01566 -0.0666 0.1700 1.0000
9.000 1.3228 0.02635 0.01686 -0.0646 0.1388 1.0000
9.250 1.3330 0.02773 0.01816 -0.0627 0.1204 1.0000
9.500 1.3444 0.02897 0.01948 -0.0608 0.1076 1.0000
9.750 1.3550 0.03020 0.02080 -0.0589 0.0927 1.0000
10.000 1.3614 0.03159 0.02217 -0.0565 0.0658 1.0000
10.250 1.3563 0.03388 0.02415 -0.0533 0.0353 1.0000
10.500 1.3534 0.03613 0.02637 -0.0507 0.0284 1.0000
10.750 1.3522 0.03839 0.02870 -0.0487 0.0258 1.0000
11.000 1.3520 0.04070 0.03118 -0.0470 0.0244 1.0000
11.250 1.3506 0.04326 0.03391 -0.0456 0.0234 1.0000
11.500 1.3476 0.04613 0.03696 -0.0446 0.0225 1.0000
11.750 1.3427 0.04938 0.04038 -0.0439 0.0216 1.0000
12.000 1.3356 0.05304 0.04425 -0.0436 0.0209 1.0000
12.250 1.3265 0.05712 0.04850 -0.0436 0.0202 1.0000
12.500 1.3169 0.06147 0.05300 -0.0440 0.0198 1.0000
12.750 1.3103 0.06559 0.05732 -0.0445 0.0193 1.0000
13.000 1.3028 0.06997 0.06189 -0.0453 0.0190 1.0000
13.250 1.2948 0.07454 0.06664 -0.0463 0.0187 1.0000
13.500 1.2867 0.07921 0.07149 -0.0474 0.0185 1.0000
13.750 1.2790 0.08394 0.07638 -0.0485 0.0182 1.0000
14.000 1.2719 0.08861 0.08121 -0.0497 0.0180 1.0000
14.250 1.2656 0.09322 0.08597 -0.0509 0.0178 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 326 (PFALZ 55) AIRFOIL (goe326-il)