GOE 178 AIRFOIL (goe178-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 178 AIRFOIL (goe178-il) Reynolds number: 1,000,000 Max Cl/Cd: 122.98 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe178-il-1000000.txt Download as CSV file: xf-goe178-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 178 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4164 0.10817 0.10661 -0.0070 1.0000 0.0141
-9.250 -0.4121 0.10409 0.10253 -0.0089 1.0000 0.0142
-8.750 -0.4075 0.09552 0.09399 -0.0123 1.0000 0.0146
-8.500 -0.3978 0.09328 0.09175 -0.0131 1.0000 0.0148
-8.250 -0.3889 0.09083 0.08931 -0.0143 1.0000 0.0151
-8.000 -0.3811 0.08807 0.08657 -0.0159 1.0000 0.0155
-7.750 -0.3746 0.08504 0.08356 -0.0179 1.0000 0.0161
-6.750 -0.3094 0.02194 0.01906 -0.0862 0.9727 0.0191
-6.500 -0.2834 0.02223 0.01942 -0.0858 0.9642 0.0196
-6.250 -0.2583 0.02344 0.02076 -0.0847 0.9546 0.0200
-6.000 -0.2343 0.02338 0.02067 -0.0841 0.9427 0.0205
-5.750 -0.2103 0.01993 0.01680 -0.0849 0.9288 0.0216
-5.500 -0.1855 0.01790 0.01443 -0.0849 0.9100 0.0225
-5.250 -0.1604 0.01736 0.01365 -0.0842 0.8796 0.0233
-5.000 -0.1352 0.01572 0.01150 -0.0840 0.8459 0.0244
-4.750 -0.1093 0.01369 0.00910 -0.0842 0.8194 0.0253
-4.250 -0.0545 0.01258 0.00765 -0.0842 0.7728 0.0269
-4.000 -0.0267 0.01210 0.00700 -0.0842 0.7473 0.0277
-3.750 0.0014 0.01148 0.00613 -0.0842 0.7153 0.0280
-3.500 0.0293 0.01096 0.00535 -0.0842 0.6745 0.0282
-3.250 0.0573 0.01058 0.00472 -0.0842 0.6322 0.0286
-3.000 0.0854 0.01026 0.00418 -0.0842 0.5952 0.0290
-2.750 0.1137 0.01006 0.00381 -0.0842 0.5649 0.0298
-2.500 0.1421 0.00982 0.00341 -0.0843 0.5405 0.0303
-2.250 0.1706 0.00963 0.00310 -0.0844 0.5208 0.0307
-2.000 0.1991 0.00949 0.00285 -0.0844 0.5046 0.0310
-1.750 0.2277 0.00935 0.00263 -0.0845 0.4913 0.0312
-1.500 0.2564 0.00914 0.00233 -0.0846 0.4800 0.0314
-1.000 0.3140 0.00878 0.00181 -0.0849 0.4629 0.0324
-0.750 0.3427 0.00868 0.00164 -0.0849 0.4560 0.0339
-0.500 0.3715 0.00859 0.00153 -0.0850 0.4505 0.0354
-0.250 0.4002 0.00854 0.00144 -0.0851 0.4448 0.0375
0.000 0.4288 0.00853 0.00139 -0.0852 0.4395 0.0400
0.250 0.4576 0.00844 0.00138 -0.0853 0.4353 0.0622
0.500 0.4862 0.00843 0.00140 -0.0854 0.4305 0.0798
0.750 0.5147 0.00846 0.00142 -0.0855 0.4259 0.0878
1.000 0.5433 0.00845 0.00144 -0.0856 0.4213 0.0968
1.250 0.5718 0.00845 0.00145 -0.0857 0.4154 0.1077
1.500 0.6003 0.00837 0.00151 -0.0859 0.4106 0.1719
1.750 0.6292 0.00802 0.00160 -0.0863 0.4072 0.3502
2.750 0.7370 0.00686 0.00189 -0.0856 0.3914 1.0000
3.000 0.7655 0.00694 0.00195 -0.0857 0.3860 1.0000
3.250 0.7936 0.00706 0.00203 -0.0857 0.3808 1.0000
3.500 0.8221 0.00713 0.00210 -0.0858 0.3766 1.0000
3.750 0.8504 0.00722 0.00217 -0.0859 0.3702 1.0000
4.000 0.8785 0.00732 0.00225 -0.0859 0.3611 1.0000
4.250 0.9065 0.00745 0.00232 -0.0860 0.3476 1.0000
4.500 0.9343 0.00760 0.00242 -0.0860 0.3340 1.0000
4.750 0.9617 0.00782 0.00255 -0.0860 0.3103 1.0000
5.000 0.9877 0.00828 0.00278 -0.0859 0.2679 1.0000
5.250 1.0133 0.00880 0.00310 -0.0857 0.2308 1.0000
5.500 1.0393 0.00924 0.00341 -0.0855 0.2051 1.0000
5.750 1.0645 0.00981 0.00377 -0.0853 0.1660 1.0000
6.000 1.0838 0.01139 0.00478 -0.0845 0.0565 1.0000
6.250 1.1097 0.01177 0.00513 -0.0843 0.0479 1.0000
6.500 1.1354 0.01219 0.00549 -0.0840 0.0311 1.0000
6.750 1.1597 0.01280 0.00599 -0.0836 0.0169 1.0000
7.000 1.1851 0.01322 0.00645 -0.0832 0.0149 1.0000
7.250 1.2103 0.01363 0.00691 -0.0829 0.0138 1.0000
7.500 1.2349 0.01412 0.00745 -0.0825 0.0128 1.0000
7.750 1.2584 0.01474 0.00813 -0.0819 0.0119 1.0000
8.000 1.2797 0.01565 0.00915 -0.0810 0.0110 1.0000
8.250 1.3034 0.01616 0.00970 -0.0805 0.0107 1.0000
8.500 1.3260 0.01676 0.01036 -0.0798 0.0103 1.0000
8.750 1.3480 0.01741 0.01108 -0.0791 0.0098 1.0000
9.000 1.3695 0.01806 0.01177 -0.0783 0.0093 1.0000
9.250 1.3898 0.01879 0.01255 -0.0774 0.0088 1.0000
9.500 1.4069 0.01980 0.01363 -0.0761 0.0084 1.0000
9.750 1.4137 0.02169 0.01565 -0.0734 0.0080 1.0000
10.000 1.4262 0.02291 0.01695 -0.0715 0.0078 1.0000
10.250 1.4420 0.02376 0.01788 -0.0700 0.0077 1.0000
10.500 1.4543 0.02478 0.01898 -0.0681 0.0075 1.0000
10.750 1.4615 0.02589 0.02017 -0.0654 0.0074 1.0000
11.000 1.4664 0.02718 0.02155 -0.0627 0.0072 1.0000
11.250 1.4706 0.02867 0.02313 -0.0603 0.0071 1.0000
11.500 1.4747 0.03031 0.02486 -0.0583 0.0069 1.0000
11.750 1.4776 0.03222 0.02686 -0.0566 0.0068 1.0000
12.000 1.4805 0.03430 0.02904 -0.0553 0.0067 1.0000
12.250 1.4835 0.03655 0.03139 -0.0544 0.0065 1.0000
12.500 1.4862 0.03896 0.03388 -0.0538 0.0064 1.0000
12.750 1.4882 0.04156 0.03658 -0.0534 0.0063 1.0000
13.000 1.4894 0.04435 0.03946 -0.0532 0.0062 1.0000
13.250 1.4907 0.04722 0.04241 -0.0533 0.0061 1.0000
13.500 1.4908 0.05031 0.04559 -0.0535 0.0060 1.0000
13.750 1.4893 0.05360 0.04896 -0.0538 0.0059 1.0000
14.000 1.4850 0.05719 0.05263 -0.0537 0.0057 1.0000
14.250 1.4747 0.06128 0.05684 -0.0527 0.0056 1.0000
14.500 1.4609 0.06599 0.06173 -0.0517 0.0055 1.0000
14.750 1.4568 0.06999 0.06586 -0.0530 0.0054 1.0000
15.000 1.4515 0.07422 0.07023 -0.0542 0.0054 1.0000
15.250 1.4451 0.07869 0.07484 -0.0555 0.0053 1.0000
15.500 1.4375 0.08344 0.07972 -0.0569 0.0053 1.0000
15.750 1.4284 0.08849 0.08491 -0.0585 0.0053 1.0000
16.000 1.4186 0.09379 0.09036 -0.0603 0.0053 1.0000
16.250 1.4079 0.09939 0.09609 -0.0624 0.0052 1.0000
16.500 1.3962 0.10529 0.10214 -0.0648 0.0052 1.0000
16.750 1.3834 0.11150 0.10849 -0.0674 0.0052 1.0000
17.000 1.3693 0.11816 0.11530 -0.0704 0.0052 1.0000
17.250 1.3556 0.12498 0.12226 -0.0737 0.0052 1.0000
17.500 1.3401 0.13236 0.12979 -0.0774 0.0052 1.0000
17.750 1.3247 0.14000 0.13757 -0.0816 0.0052 1.0000
18.000 1.3082 0.14824 0.14596 -0.0863 0.0053 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 178 AIRFOIL (goe178-il)