GOE 432 AIRFOIL (goe432-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 432 AIRFOIL (goe432-il) Reynolds number: 1,000,000 Max Cl/Cd: 107.98 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe432-il-1000000.txt Download as CSV file: xf-goe432-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 432 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3711 0.09598 0.09438 -0.0197 1.0000 0.0113
-8.500 -0.3719 0.09341 0.09183 -0.0194 1.0000 0.0116
-7.250 -0.4023 0.01877 0.01552 -0.0787 0.9848 0.0104
-7.000 -0.3699 0.01708 0.01359 -0.0806 0.9838 0.0108
-6.750 -0.3449 0.01591 0.01224 -0.0803 0.9804 0.0112
-6.500 -0.3154 0.01475 0.01088 -0.0809 0.9777 0.0117
-6.250 -0.2832 0.01365 0.00957 -0.0821 0.9757 0.0122
-6.000 -0.2482 0.01272 0.00845 -0.0836 0.9740 0.0127
-5.750 -0.2094 0.01199 0.00756 -0.0859 0.9726 0.0131
-5.500 -0.1724 0.01057 0.00593 -0.0880 0.9713 0.0143
-5.250 -0.1426 0.01019 0.00552 -0.0883 0.9676 0.0152
-5.000 -0.1132 0.00987 0.00514 -0.0884 0.9635 0.0162
-4.750 -0.0812 0.00951 0.00471 -0.0891 0.9605 0.0170
-4.500 -0.0486 0.00886 0.00401 -0.0899 0.9570 0.0189
-4.250 -0.0226 0.00863 0.00376 -0.0892 0.9491 0.0203
-4.000 0.0084 0.00841 0.00348 -0.0896 0.9421 0.0216
-3.750 0.0333 0.00805 0.00312 -0.0887 0.9329 0.0243
-3.500 0.0620 0.00797 0.00304 -0.0886 0.9232 0.0266
-3.250 0.0892 0.00766 0.00267 -0.0882 0.9097 0.0291
-3.000 0.1170 0.00756 0.00255 -0.0879 0.8906 0.0317
-2.750 0.1447 0.00756 0.00245 -0.0875 0.8637 0.0341
-2.500 0.1707 0.00756 0.00231 -0.0868 0.8313 0.0358
-2.250 0.1949 0.00738 0.00197 -0.0858 0.7992 0.0390
-2.000 0.2194 0.00745 0.00190 -0.0848 0.7650 0.0415
-1.750 0.2434 0.00754 0.00182 -0.0838 0.7234 0.0437
-1.500 0.2674 0.00770 0.00179 -0.0827 0.6775 0.0450
-1.250 0.2906 0.00764 0.00149 -0.0816 0.6267 0.0480
-1.000 0.3145 0.00773 0.00139 -0.0807 0.5782 0.0515
-0.750 0.3397 0.00780 0.00132 -0.0800 0.5473 0.0544
-0.500 0.3656 0.00786 0.00127 -0.0794 0.5263 0.0568
-0.250 0.3918 0.00785 0.00119 -0.0789 0.5099 0.0614
0.000 0.4181 0.00784 0.00115 -0.0785 0.4953 0.0701
0.250 0.4439 0.00771 0.00113 -0.0780 0.4815 0.1276
0.500 0.4687 0.00743 0.00119 -0.0774 0.4682 0.2788
0.750 0.4943 0.00729 0.00123 -0.0770 0.4546 0.3648
1.000 0.5113 0.00636 0.00133 -0.0749 0.4424 0.7645
1.250 0.5827 0.00609 0.00142 -0.0846 0.4228 1.0000
1.500 0.6079 0.00620 0.00145 -0.0839 0.4096 1.0000
1.750 0.6330 0.00633 0.00150 -0.0832 0.3952 1.0000
2.250 0.6829 0.00663 0.00162 -0.0817 0.3628 1.0000
2.500 0.7079 0.00678 0.00169 -0.0810 0.3491 1.0000
2.750 0.7328 0.00694 0.00178 -0.0803 0.3350 1.0000
3.000 0.7577 0.00711 0.00187 -0.0796 0.3195 1.0000
3.250 0.7825 0.00728 0.00197 -0.0789 0.3040 1.0000
3.500 0.8070 0.00749 0.00208 -0.0782 0.2824 1.0000
3.750 0.8302 0.00782 0.00224 -0.0773 0.2471 1.0000
4.000 0.8518 0.00832 0.00248 -0.0761 0.1973 1.0000
4.250 0.8757 0.00861 0.00268 -0.0753 0.1834 1.0000
4.500 0.9004 0.00882 0.00286 -0.0746 0.1767 1.0000
4.750 0.9255 0.00900 0.00303 -0.0740 0.1732 1.0000
5.000 0.9504 0.00920 0.00322 -0.0734 0.1698 1.0000
5.250 0.9751 0.00942 0.00342 -0.0727 0.1660 1.0000
5.500 0.9999 0.00963 0.00364 -0.0721 0.1631 1.0000
5.750 1.0252 0.00978 0.00382 -0.0715 0.1616 1.0000
6.000 1.0503 0.00994 0.00402 -0.0710 0.1596 1.0000
6.250 1.0752 0.01014 0.00423 -0.0704 0.1568 1.0000
6.500 1.0993 0.01041 0.00449 -0.0697 0.1517 1.0000
6.750 1.1239 0.01062 0.00471 -0.0691 0.1478 1.0000
7.000 1.1495 0.01073 0.00485 -0.0687 0.1434 1.0000
7.250 1.1738 0.01096 0.00507 -0.0681 0.1370 1.0000
7.500 1.1988 0.01114 0.00527 -0.0675 0.1320 1.0000
7.750 1.2234 0.01133 0.00546 -0.0670 0.1250 1.0000
8.000 1.2474 0.01158 0.00569 -0.0663 0.1153 1.0000
8.250 1.2694 0.01201 0.00599 -0.0654 0.0955 1.0000
8.500 1.2908 0.01249 0.00640 -0.0644 0.0824 1.0000
8.750 1.3119 0.01299 0.00686 -0.0633 0.0717 1.0000
9.000 1.3304 0.01371 0.00745 -0.0619 0.0495 1.0000
9.250 1.3438 0.01489 0.00843 -0.0597 0.0228 1.0000
9.500 1.3620 0.01560 0.00914 -0.0581 0.0171 1.0000
9.750 1.3800 0.01631 0.00987 -0.0566 0.0142 1.0000
10.000 1.3998 0.01681 0.01044 -0.0553 0.0131 1.0000
10.250 1.4179 0.01743 0.01110 -0.0538 0.0120 1.0000
10.500 1.4319 0.01838 0.01213 -0.0517 0.0107 1.0000
10.750 1.4495 0.01895 0.01279 -0.0501 0.0103 1.0000
11.000 1.4660 0.01957 0.01348 -0.0484 0.0098 1.0000
11.250 1.4807 0.02026 0.01424 -0.0465 0.0093 1.0000
11.500 1.4926 0.02095 0.01499 -0.0440 0.0088 1.0000
11.750 1.5008 0.02179 0.01590 -0.0410 0.0084 1.0000
12.000 1.5043 0.02292 0.01712 -0.0375 0.0080 1.0000
12.250 1.5050 0.02424 0.01856 -0.0338 0.0076 1.0000
12.500 1.5126 0.02518 0.01961 -0.0313 0.0075 1.0000
12.750 1.5187 0.02626 0.02077 -0.0288 0.0073 1.0000
13.000 1.5242 0.02744 0.02204 -0.0265 0.0071 1.0000
13.250 1.5276 0.02882 0.02352 -0.0242 0.0069 1.0000
13.500 1.5300 0.03036 0.02516 -0.0221 0.0067 1.0000
13.750 1.5322 0.03204 0.02693 -0.0204 0.0065 1.0000
14.000 1.5300 0.03426 0.02926 -0.0188 0.0064 1.0000
14.250 1.5281 0.03666 0.03177 -0.0177 0.0062 1.0000
14.500 1.5255 0.03935 0.03455 -0.0172 0.0061 1.0000
14.750 1.5199 0.04266 0.03798 -0.0171 0.0060 1.0000
15.000 1.5119 0.04657 0.04201 -0.0177 0.0059 1.0000
15.250 1.4966 0.05188 0.04746 -0.0193 0.0058 1.0000
15.500 1.4846 0.05707 0.05279 -0.0211 0.0058 1.0000
15.750 1.4727 0.06243 0.05828 -0.0233 0.0057 1.0000
16.000 1.4550 0.06877 0.06477 -0.0258 0.0057 1.0000
16.250 1.4376 0.07516 0.07128 -0.0285 0.0057 1.0000
16.500 1.4206 0.08169 0.07793 -0.0313 0.0057 1.0000
16.750 1.4039 0.08830 0.08467 -0.0342 0.0056 1.0000
17.000 1.3886 0.09492 0.09139 -0.0372 0.0057 1.0000
17.250 1.3644 0.10317 0.09977 -0.0411 0.0055 1.0000
17.500 1.3493 0.11005 0.10676 -0.0444 0.0055 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 432 AIRFOIL (goe432-il)