GOE 377 AIRFOIL (goe377-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 377 AIRFOIL (goe377-il) Reynolds number: 1,000,000 Max Cl/Cd: 105.16 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe377-il-1000000-n5.txt Download as CSV file: xf-goe377-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 377 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.2832 0.10299 0.10145 -0.0267 1.0000 0.0044
-10.000 -0.3827 0.11234 0.11067 -0.0203 1.0000 0.0030
-9.750 -0.3801 0.10826 0.10660 -0.0214 1.0000 0.0036
-9.500 -0.3762 0.10542 0.10377 -0.0220 1.0000 0.0034
-9.250 -0.3719 0.10277 0.10115 -0.0227 1.0000 0.0041
-8.750 -0.3680 0.09642 0.09483 -0.0238 1.0000 0.0033
-8.250 -0.2696 0.07228 0.07083 -0.0338 0.9935 0.0036
-8.000 -0.3684 0.08755 0.08604 -0.0245 0.9999 0.0036
-7.750 -0.3550 0.08306 0.08155 -0.0297 0.9966 0.0035
-7.500 -0.3420 0.07875 0.07725 -0.0351 0.9890 0.0035
-7.250 -0.3245 0.07370 0.07219 -0.0420 0.9812 0.0034
-6.750 -0.2724 0.06253 0.06093 -0.0593 0.9672 0.0032
-6.250 -0.2541 0.01551 0.01185 -0.0847 0.9248 0.0036
-6.000 -0.2322 0.01343 0.00938 -0.0839 0.9162 0.0037
-5.750 -0.2093 0.01180 0.00742 -0.0831 0.9092 0.0040
-5.500 -0.1843 0.01100 0.00643 -0.0826 0.9023 0.0043
-5.250 -0.1585 0.01047 0.00578 -0.0823 0.8963 0.0046
-5.000 -0.1328 0.00997 0.00515 -0.0818 0.8894 0.0050
-4.750 -0.1067 0.00956 0.00462 -0.0814 0.8833 0.0055
-4.500 -0.0806 0.00919 0.00414 -0.0810 0.8765 0.0059
-4.250 -0.0551 0.00864 0.00344 -0.0804 0.8699 0.0068
-4.000 -0.0292 0.00830 0.00303 -0.0800 0.8625 0.0077
-3.750 -0.0031 0.00804 0.00268 -0.0795 0.8554 0.0087
-3.500 0.0231 0.00778 0.00234 -0.0791 0.8473 0.0102
-3.250 0.0493 0.00759 0.00216 -0.0786 0.8390 0.0140
-3.000 0.0758 0.00754 0.00212 -0.0783 0.8283 0.0195
-2.750 0.1022 0.00757 0.00210 -0.0779 0.8122 0.0233
-2.500 0.1281 0.00770 0.00213 -0.0773 0.7887 0.0254
-2.250 0.1534 0.00780 0.00212 -0.0767 0.7613 0.0260
-2.000 0.1786 0.00792 0.00213 -0.0760 0.7332 0.0264
-1.750 0.2030 0.00777 0.00183 -0.0752 0.7084 0.0271
-1.250 0.2527 0.00751 0.00134 -0.0739 0.6667 0.0292
-1.000 0.2779 0.00746 0.00118 -0.0733 0.6464 0.0294
-0.750 0.3033 0.00743 0.00104 -0.0727 0.6255 0.0298
-0.500 0.3285 0.00744 0.00094 -0.0721 0.6022 0.0305
-0.250 0.3541 0.00745 0.00087 -0.0716 0.5832 0.0322
0.000 0.3799 0.00747 0.00083 -0.0711 0.5655 0.0351
0.250 0.4058 0.00751 0.00081 -0.0707 0.5487 0.0372
0.500 0.4315 0.00754 0.00080 -0.0703 0.5319 0.0437
1.250 0.5074 0.00772 0.00088 -0.0688 0.4667 0.0994
1.500 0.5322 0.00764 0.00095 -0.0683 0.4496 0.1910
1.750 0.5554 0.00742 0.00105 -0.0675 0.4306 0.3628
2.250 0.6658 0.00652 0.00139 -0.0809 0.3608 1.0000
2.500 0.6892 0.00674 0.00149 -0.0800 0.3365 1.0000
2.750 0.7122 0.00699 0.00161 -0.0791 0.3096 1.0000
3.000 0.7359 0.00720 0.00173 -0.0783 0.2900 1.0000
3.500 0.7832 0.00760 0.00198 -0.0767 0.2559 1.0000
3.750 0.8076 0.00775 0.00211 -0.0760 0.2462 1.0000
4.000 0.8318 0.00791 0.00224 -0.0753 0.2360 1.0000
4.250 0.8546 0.00819 0.00242 -0.0744 0.2140 1.0000
4.500 0.8776 0.00845 0.00260 -0.0736 0.1930 1.0000
4.750 0.8983 0.00892 0.00286 -0.0723 0.1508 1.0000
5.000 0.9168 0.00959 0.00326 -0.0707 0.0998 1.0000
5.250 0.9355 0.01026 0.00370 -0.0691 0.0513 1.0000
5.500 0.9559 0.01080 0.00410 -0.0678 0.0229 1.0000
5.750 0.9780 0.01117 0.00444 -0.0668 0.0119 1.0000
6.000 1.0003 0.01152 0.00483 -0.0657 0.0084 1.0000
6.250 1.0232 0.01180 0.00514 -0.0649 0.0074 1.0000
6.500 1.0453 0.01215 0.00550 -0.0639 0.0062 1.0000
6.750 1.0667 0.01258 0.00598 -0.0628 0.0052 1.0000
7.000 1.0887 0.01293 0.00638 -0.0617 0.0048 1.0000
7.250 1.1102 0.01333 0.00684 -0.0607 0.0043 1.0000
7.500 1.1314 0.01375 0.00730 -0.0595 0.0040 1.0000
7.750 1.1511 0.01431 0.00790 -0.0582 0.0036 1.0000
8.000 1.1708 0.01484 0.00850 -0.0569 0.0032 1.0000
8.250 1.1907 0.01535 0.00907 -0.0555 0.0029 1.0000
8.500 1.2102 0.01587 0.00966 -0.0542 0.0028 1.0000
8.750 1.2284 0.01649 0.01035 -0.0527 0.0026 1.0000
9.000 1.2460 0.01713 0.01107 -0.0511 0.0024 1.0000
9.250 1.2641 0.01770 0.01171 -0.0496 0.0023 1.0000
9.500 1.2797 0.01845 0.01253 -0.0478 0.0021 1.0000
9.750 1.2913 0.01948 0.01367 -0.0452 0.0020 1.0000
10.000 1.3020 0.02050 0.01481 -0.0426 0.0019 1.0000
10.250 1.3118 0.02144 0.01585 -0.0399 0.0019 1.0000
10.500 1.3208 0.02227 0.01679 -0.0369 0.0018 1.0000
10.750 1.3278 0.02327 0.01791 -0.0338 0.0017 1.0000
11.000 1.3303 0.02464 0.01941 -0.0303 0.0017 1.0000
11.250 1.3351 0.02590 0.02081 -0.0274 0.0017 1.0000
11.500 1.3380 0.02739 0.02245 -0.0245 0.0016 1.0000
11.750 1.3401 0.02901 0.02421 -0.0218 0.0016 1.0000
12.000 1.3396 0.03096 0.02631 -0.0191 0.0014 1.0000
12.250 1.3399 0.03290 0.02840 -0.0169 0.0015 1.0000
12.500 1.3502 0.03387 0.02943 -0.0159 0.0013 1.0000
12.750 1.3478 0.03624 0.03196 -0.0141 0.0013 1.0000
13.000 1.3392 0.03943 0.03535 -0.0123 0.0013 1.0000
13.250 1.3405 0.04168 0.03771 -0.0117 0.0013 1.0000
13.500 1.3358 0.04478 0.04096 -0.0112 0.0013 1.0000
13.750 1.3355 0.04752 0.04380 -0.0112 0.0012 1.0000
14.000 1.3223 0.05215 0.04862 -0.0117 0.0012 1.0000
14.250 1.3001 0.05847 0.05519 -0.0130 0.0012 1.0000
14.500 1.3048 0.06108 0.05789 -0.0142 0.0012 1.0000
14.750 1.3048 0.06457 0.06145 -0.0157 0.0012 1.0000
15.000 1.3002 0.06897 0.06595 -0.0177 0.0011 1.0000
15.250 1.2711 0.07770 0.07490 -0.0214 0.0011 1.0000
15.500 1.2609 0.08349 0.08081 -0.0241 0.0012 1.0000
15.750 1.2307 0.09332 0.09085 -0.0287 0.0012 1.0000
16.000 1.2175 0.10021 0.09784 -0.0322 0.0012 1.0000
16.250 1.1886 0.11062 0.10844 -0.0374 0.0012 1.0000
16.500 1.1721 0.11868 0.11660 -0.0416 0.0012 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 377 AIRFOIL (goe377-il)