GOE 769 AIRFOIL (goe769-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 769 AIRFOIL (goe769-il) Reynolds number: 200,000 Max Cl/Cd: 59.05 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe769-il-200000.txt Download as CSV file: xf-goe769-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 769 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2179 0.10426 0.10075 -0.0219 1.0000 0.0654
-9.000 -0.2185 0.10151 0.09805 -0.0238 1.0000 0.0676
-8.750 -0.2368 0.09800 0.09463 -0.0286 1.0000 0.0688
-8.500 -0.2196 0.09534 0.09202 -0.0267 1.0000 0.0693
-8.250 -0.2087 0.09300 0.08975 -0.0258 1.0000 0.0700
-8.000 -0.2027 0.09075 0.08758 -0.0256 1.0000 0.0709
-7.750 -0.1887 0.08755 0.08446 -0.0286 0.9837 0.0726
-7.500 -0.1716 0.07938 0.07620 -0.0494 0.9177 0.0762
-7.250 -0.1234 0.07587 0.07257 -0.0518 0.8815 0.0772
-7.000 -0.0999 0.07308 0.06960 -0.0543 0.8410 0.0786
-6.750 -0.0957 0.07074 0.06713 -0.0549 0.8107 0.0801
-6.500 -0.1281 0.06407 0.06021 -0.0677 0.7890 0.0842
-6.250 -0.1118 0.06225 0.05835 -0.0652 0.7723 0.0847
-6.000 -0.0966 0.06067 0.05670 -0.0632 0.7567 0.0855
-5.750 -0.0825 0.05895 0.05494 -0.0624 0.7414 0.0870
-5.500 -0.0827 0.05256 0.04812 -0.0710 0.7313 0.0937
-5.250 -0.0652 0.05074 0.04632 -0.0695 0.7176 0.0944
-5.000 -0.0480 0.04900 0.04455 -0.0686 0.7048 0.0954
-4.750 -0.0536 0.03223 0.02637 -0.0742 0.7014 0.0690
-4.500 -0.0324 0.03011 0.02405 -0.0735 0.6904 0.0673
-4.250 -0.0126 0.02719 0.02076 -0.0728 0.6789 0.0659
-4.000 0.0088 0.02467 0.01773 -0.0719 0.6684 0.0661
-3.750 0.0320 0.02278 0.01542 -0.0709 0.6565 0.0668
-3.500 0.0565 0.02125 0.01349 -0.0700 0.6448 0.0675
-3.250 0.0820 0.02009 0.01189 -0.0690 0.6333 0.0684
-3.000 0.1078 0.01919 0.01095 -0.0684 0.6206 0.0698
-2.750 0.1341 0.01868 0.01032 -0.0676 0.6087 0.0716
-2.500 0.1605 0.01813 0.00960 -0.0668 0.5947 0.0746
-2.250 0.1868 0.01758 0.00897 -0.0661 0.5813 0.0777
-2.000 0.2134 0.01723 0.00850 -0.0654 0.5693 0.0813
-1.750 0.2400 0.01668 0.00789 -0.0647 0.5560 0.0855
-1.500 0.2667 0.01642 0.00758 -0.0640 0.5442 0.0912
-1.250 0.2931 0.01611 0.00726 -0.0633 0.5327 0.0985
-1.000 0.3196 0.01580 0.00696 -0.0627 0.5208 0.1079
-0.750 0.3460 0.01558 0.00665 -0.0620 0.5104 0.1190
-0.500 0.3725 0.01535 0.00645 -0.0614 0.4987 0.1313
-0.250 0.3991 0.01527 0.00630 -0.0608 0.4885 0.1442
0.000 0.4252 0.01508 0.00614 -0.0602 0.4780 0.1583
0.250 0.4512 0.01492 0.00602 -0.0595 0.4682 0.1739
0.500 0.4772 0.01478 0.00589 -0.0589 0.4584 0.1919
0.750 0.5032 0.01468 0.00584 -0.0582 0.4491 0.2153
1.000 0.5284 0.01445 0.00580 -0.0575 0.4400 0.2610
1.250 0.5457 0.01340 0.00587 -0.0556 0.4322 0.5764
1.500 0.6197 0.01280 0.00586 -0.0638 0.4185 1.0000
1.750 0.6444 0.01298 0.00595 -0.0629 0.4094 1.0000
2.000 0.6690 0.01321 0.00600 -0.0620 0.4011 1.0000
2.250 0.6937 0.01344 0.00614 -0.0611 0.3923 1.0000
2.500 0.7183 0.01365 0.00622 -0.0602 0.3836 1.0000
2.750 0.7429 0.01391 0.00638 -0.0594 0.3750 1.0000
3.000 0.7675 0.01413 0.00650 -0.0585 0.3663 1.0000
3.250 0.7922 0.01443 0.00668 -0.0577 0.3583 1.0000
3.500 0.8167 0.01465 0.00685 -0.0568 0.3496 1.0000
3.750 0.8412 0.01499 0.00704 -0.0560 0.3418 1.0000
4.000 0.8655 0.01521 0.00725 -0.0552 0.3332 1.0000
4.250 0.8900 0.01560 0.00745 -0.0544 0.3261 1.0000
4.500 0.9141 0.01585 0.00774 -0.0536 0.3184 1.0000
4.750 0.9384 0.01615 0.00794 -0.0528 0.3115 1.0000
5.000 0.9626 0.01653 0.00826 -0.0520 0.3048 1.0000
5.250 0.9866 0.01682 0.00853 -0.0512 0.2980 1.0000
5.500 1.0108 0.01724 0.00880 -0.0505 0.2924 1.0000
5.750 1.0345 0.01759 0.00919 -0.0497 0.2865 1.0000
6.000 1.0582 0.01792 0.00949 -0.0489 0.2808 1.0000
6.250 1.0822 0.01836 0.00980 -0.0482 0.2758 1.0000
6.500 1.1054 0.01876 0.01025 -0.0473 0.2707 1.0000
6.750 1.1287 0.01914 0.01063 -0.0465 0.2658 1.0000
7.000 1.1521 0.01955 0.01097 -0.0458 0.2615 1.0000
7.250 1.1756 0.02011 0.01146 -0.0451 0.2572 1.0000
7.500 1.1975 0.02051 0.01194 -0.0441 0.2527 1.0000
7.750 1.2200 0.02093 0.01235 -0.0433 0.2485 1.0000
8.000 1.2432 0.02143 0.01278 -0.0426 0.2449 1.0000
8.250 1.2661 0.02208 0.01341 -0.0419 0.2413 1.0000
8.500 1.2866 0.02258 0.01400 -0.0409 0.2376 1.0000
8.750 1.3077 0.02306 0.01450 -0.0399 0.2339 1.0000
9.000 1.3297 0.02357 0.01496 -0.0391 0.2305 1.0000
9.250 1.3544 0.02439 0.01567 -0.0388 0.2271 1.0000
9.500 1.3718 0.02496 0.01640 -0.0374 0.2243 1.0000
9.750 1.3903 0.02558 0.01710 -0.0362 0.2212 1.0000
10.000 1.4099 0.02616 0.01771 -0.0352 0.2181 1.0000
10.250 1.4310 0.02676 0.01827 -0.0344 0.2153 1.0000
10.500 1.4566 0.02773 0.01912 -0.0344 0.2121 1.0000
10.750 1.4688 0.02842 0.02000 -0.0325 0.2098 1.0000
11.000 1.4826 0.02918 0.02089 -0.0308 0.2072 1.0000
11.250 1.4980 0.02992 0.02171 -0.0294 0.2046 1.0000
11.500 1.5154 0.03063 0.02243 -0.0282 0.2022 1.0000
11.750 1.5365 0.03135 0.02311 -0.0276 0.1999 1.0000
12.000 1.5592 0.03252 0.02424 -0.0275 0.1972 1.0000
12.250 1.5610 0.03342 0.02533 -0.0244 0.1955 1.0000
12.500 1.5632 0.03444 0.02650 -0.0215 0.1934 1.0000
12.750 1.5685 0.03551 0.02770 -0.0193 0.1914 1.0000
13.000 1.5771 0.03655 0.02881 -0.0176 0.1894 1.0000
13.250 1.5902 0.03750 0.02979 -0.0165 0.1875 1.0000
13.500 1.6098 0.03836 0.03061 -0.0160 0.1857 1.0000
13.750 1.6338 0.03972 0.03194 -0.0161 0.1834 1.0000
14.000 1.6228 0.04147 0.03392 -0.0133 0.1822 1.0000
14.250 1.6124 0.04355 0.03620 -0.0113 0.1806 1.0000
14.500 1.6026 0.04593 0.03877 -0.0098 0.1789 1.0000
14.750 1.5950 0.04844 0.04144 -0.0089 0.1773 1.0000
15.000 1.5907 0.05091 0.04404 -0.0083 0.1758 1.0000
15.250 1.5931 0.05296 0.04617 -0.0080 0.1742 1.0000
15.500 1.6048 0.05437 0.04759 -0.0076 0.1727 1.0000
15.750 1.6296 0.05503 0.04819 -0.0073 0.1712 1.0000
16.000 1.6184 0.05870 0.05202 -0.0072 0.1698 1.0000
16.250 1.5567 0.06706 0.06076 -0.0090 0.1687 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 769 AIRFOIL (goe769-il)