GOE 693 AIRFOIL (goe693-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 693 AIRFOIL (goe693-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.54 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe693-il-1000000.txt Download as CSV file: xf-goe693-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 693 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
------ -------- --------- --------- -------- -------- --------
-15.250 -1.1162 0.03654 0.03357 -0.0903 1.0000 0.0168
-15.000 -1.1298 0.03310 0.02998 -0.0907 1.0000 0.0170
-14.750 -1.1318 0.03134 0.02813 -0.0888 1.0000 0.0172
-14.500 -1.1232 0.03010 0.02683 -0.0876 1.0000 0.0175
-14.250 -1.1136 0.02908 0.02575 -0.0860 1.0000 0.0178
-14.000 -1.1057 0.02818 0.02477 -0.0838 1.0000 0.0181
-13.750 -1.0989 0.02733 0.02386 -0.0812 1.0000 0.0185
-13.500 -1.0919 0.02654 0.02298 -0.0784 1.0000 0.0188
-13.250 -1.0835 0.02578 0.02212 -0.0758 1.0000 0.0191
-13.000 -1.0639 0.02499 0.02123 -0.0752 0.9994 0.0195
-12.750 -1.0347 0.02431 0.02044 -0.0764 0.9982 0.0198
-12.500 -1.0126 0.02238 0.01836 -0.0774 0.9964 0.0206
-12.250 -0.9823 0.02173 0.01769 -0.0787 0.9952 0.0211
-12.000 -0.9507 0.02128 0.01720 -0.0801 0.9943 0.0217
-11.750 -0.9201 0.02075 0.01662 -0.0812 0.9930 0.0223
-11.500 -0.8911 0.02017 0.01595 -0.0821 0.9910 0.0229
-11.250 -0.8607 0.01960 0.01529 -0.0831 0.9892 0.0234
-11.000 -0.8289 0.01917 0.01477 -0.0843 0.9877 0.0238
-10.750 -0.8016 0.01767 0.01313 -0.0854 0.9861 0.0247
-10.500 -0.7695 0.01710 0.01254 -0.0868 0.9848 0.0254
-10.250 -0.7372 0.01667 0.01208 -0.0881 0.9833 0.0261
-10.000 -0.7094 0.01623 0.01158 -0.0884 0.9798 0.0267
-9.750 -0.6816 0.01576 0.01104 -0.0886 0.9759 0.0274
-9.500 -0.6533 0.01537 0.01058 -0.0889 0.9722 0.0280
-9.250 -0.6259 0.01510 0.01025 -0.0888 0.9680 0.0284
-9.000 -0.6038 0.01428 0.00932 -0.0880 0.9617 0.0291
-8.750 -0.5804 0.01357 0.00856 -0.0873 0.9565 0.0300
-8.500 -0.5557 0.01319 0.00814 -0.0867 0.9503 0.0307
-8.250 -0.5306 0.01287 0.00779 -0.0861 0.9439 0.0314
-8.000 -0.5052 0.01253 0.00739 -0.0856 0.9375 0.0322
-7.750 -0.4796 0.01218 0.00698 -0.0851 0.9297 0.0328
-7.500 -0.4536 0.01190 0.00663 -0.0847 0.9224 0.0335
-7.250 -0.4269 0.01166 0.00634 -0.0844 0.9135 0.0340
-7.000 -0.4018 0.01107 0.00566 -0.0839 0.9043 0.0350
-6.750 -0.3760 0.01066 0.00518 -0.0834 0.8939 0.0361
-6.500 -0.3492 0.01037 0.00485 -0.0832 0.8818 0.0371
-6.250 -0.3221 0.01014 0.00455 -0.0830 0.8693 0.0381
-6.000 -0.2948 0.00995 0.00428 -0.0828 0.8567 0.0393
-5.750 -0.2674 0.00979 0.00403 -0.0826 0.8439 0.0402
-5.500 -0.2404 0.00950 0.00366 -0.0824 0.8310 0.0418
-5.250 -0.2130 0.00926 0.00338 -0.0822 0.8183 0.0440
-5.000 -0.1852 0.00910 0.00316 -0.0822 0.8068 0.0461
-4.750 -0.1572 0.00899 0.00298 -0.0821 0.7963 0.0481
-4.500 -0.1297 0.00876 0.00273 -0.0820 0.7854 0.0536
-4.250 -0.1017 0.00860 0.00255 -0.0820 0.7749 0.0600
-4.000 -0.0738 0.00846 0.00241 -0.0820 0.7650 0.0693
-3.750 -0.0455 0.00835 0.00231 -0.0821 0.7559 0.0794
-3.500 -0.0171 0.00828 0.00225 -0.0821 0.7482 0.0879
-3.250 0.0114 0.00822 0.00216 -0.0822 0.7401 0.0942
-3.000 0.0399 0.00817 0.00210 -0.0823 0.7327 0.1001
-2.750 0.0687 0.00816 0.00205 -0.0824 0.7253 0.1034
-2.500 0.0970 0.00806 0.00195 -0.0825 0.7181 0.1100
-2.250 0.1256 0.00800 0.00188 -0.0826 0.7096 0.1146
-2.000 0.1542 0.00799 0.00182 -0.0827 0.7013 0.1176
-1.750 0.1826 0.00791 0.00173 -0.0828 0.6929 0.1241
-1.500 0.2111 0.00786 0.00168 -0.0829 0.6858 0.1303
-1.250 0.2399 0.00780 0.00162 -0.0830 0.6789 0.1385
-1.000 0.2683 0.00775 0.00158 -0.0831 0.6729 0.1489
-0.750 0.2969 0.00763 0.00154 -0.0833 0.6675 0.1699
-0.500 0.3253 0.00748 0.00151 -0.0834 0.6617 0.2082
-0.250 0.3533 0.00734 0.00150 -0.0835 0.6560 0.2601
0.000 0.3815 0.00711 0.00149 -0.0837 0.6503 0.3298
0.250 0.4089 0.00683 0.00149 -0.0838 0.6439 0.4345
0.500 0.4355 0.00643 0.00152 -0.0837 0.6375 0.5823
0.750 0.4608 0.00601 0.00156 -0.0832 0.6293 0.7400
1.000 0.4827 0.00566 0.00163 -0.0817 0.6217 0.8816
1.250 0.5128 0.00561 0.00170 -0.0817 0.6135 0.9674
1.500 0.5518 0.00569 0.00174 -0.0841 0.6052 0.9867
1.750 0.5912 0.00578 0.00178 -0.0867 0.5954 0.9948
2.000 0.6320 0.00588 0.00182 -0.0895 0.5820 0.9993
2.250 0.6605 0.00597 0.00185 -0.0897 0.5686 1.0000
2.500 0.6861 0.00607 0.00189 -0.0892 0.5550 1.0000
2.750 0.7113 0.00619 0.00193 -0.0887 0.5369 1.0000
3.000 0.7360 0.00637 0.00200 -0.0880 0.5141 1.0000
3.250 0.7602 0.00658 0.00209 -0.0873 0.4858 1.0000
3.500 0.7836 0.00688 0.00222 -0.0865 0.4514 1.0000
3.750 0.8070 0.00719 0.00238 -0.0857 0.4186 1.0000
4.000 0.8308 0.00750 0.00255 -0.0850 0.3911 1.0000
4.250 0.8555 0.00775 0.00270 -0.0844 0.3713 1.0000
4.500 0.8806 0.00797 0.00286 -0.0839 0.3560 1.0000
4.750 0.9059 0.00819 0.00302 -0.0835 0.3430 1.0000
5.000 0.9314 0.00841 0.00318 -0.0831 0.3306 1.0000
5.250 0.9569 0.00862 0.00334 -0.0827 0.3180 1.0000
5.500 0.9825 0.00882 0.00351 -0.0824 0.3072 1.0000
5.750 1.0079 0.00905 0.00369 -0.0820 0.2961 1.0000
6.000 1.0330 0.00929 0.00388 -0.0816 0.2819 1.0000
6.250 1.0569 0.00962 0.00410 -0.0810 0.2557 1.0000
6.500 1.0761 0.01030 0.00449 -0.0797 0.2022 1.0000
6.750 1.0971 0.01083 0.00486 -0.0786 0.1783 1.0000
7.000 1.1204 0.01118 0.00516 -0.0779 0.1690 1.0000
7.250 1.1443 0.01146 0.00544 -0.0773 0.1621 1.0000
7.500 1.1684 0.01173 0.00571 -0.0768 0.1569 1.0000
7.750 1.1913 0.01208 0.00602 -0.0760 0.1502 1.0000
8.000 1.2160 0.01228 0.00626 -0.0756 0.1454 1.0000
8.250 1.2387 0.01261 0.00655 -0.0748 0.1376 1.0000
8.500 1.2621 0.01288 0.00683 -0.0742 0.1298 1.0000
8.750 1.2822 0.01336 0.00718 -0.0730 0.1094 1.0000
9.000 1.2986 0.01405 0.00771 -0.0713 0.0874 1.0000
9.250 1.3172 0.01456 0.00818 -0.0699 0.0787 1.0000
9.500 1.3354 0.01505 0.00866 -0.0684 0.0707 1.0000
9.750 1.3481 0.01571 0.00924 -0.0660 0.0547 1.0000
10.000 1.3515 0.01689 0.01022 -0.0621 0.0303 1.0000
10.250 1.3636 0.01767 0.01098 -0.0597 0.0247 1.0000
10.500 1.3762 0.01843 0.01175 -0.0576 0.0217 1.0000
10.750 1.3910 0.01910 0.01245 -0.0558 0.0202 1.0000
11.000 1.4049 0.01983 0.01322 -0.0540 0.0189 1.0000
11.250 1.4159 0.02077 0.01419 -0.0519 0.0175 1.0000
11.500 1.4298 0.02155 0.01502 -0.0503 0.0168 1.0000
11.750 1.4434 0.02236 0.01588 -0.0487 0.0161 1.0000
12.000 1.4558 0.02328 0.01684 -0.0471 0.0155 1.0000
12.250 1.4664 0.02436 0.01796 -0.0454 0.0148 1.0000
12.500 1.4728 0.02579 0.01947 -0.0434 0.0142 1.0000
12.750 1.4819 0.02706 0.02081 -0.0418 0.0138 1.0000
13.000 1.4926 0.02824 0.02205 -0.0405 0.0135 1.0000
13.250 1.5021 0.02957 0.02345 -0.0392 0.0131 1.0000
13.500 1.5106 0.03101 0.02495 -0.0379 0.0128 1.0000
13.750 1.5184 0.03255 0.02655 -0.0367 0.0125 1.0000
14.000 1.5252 0.03423 0.02830 -0.0356 0.0121 1.0000
14.250 1.5296 0.03618 0.03031 -0.0345 0.0118 1.0000
14.500 1.5289 0.03865 0.03287 -0.0333 0.0115 1.0000
14.750 1.5210 0.04195 0.03628 -0.0320 0.0111 1.0000
15.000 1.5276 0.04388 0.03828 -0.0314 0.0110 1.0000
15.250 1.5310 0.04617 0.04066 -0.0309 0.0109 1.0000
15.500 1.5332 0.04866 0.04324 -0.0304 0.0107 1.0000
15.750 1.5338 0.05140 0.04607 -0.0300 0.0105 1.0000
16.000 1.5334 0.05433 0.04909 -0.0298 0.0103 1.0000
16.250 1.5320 0.05745 0.05230 -0.0297 0.0102 1.0000
16.500 1.5296 0.06077 0.05571 -0.0297 0.0100 1.0000
16.750 1.5262 0.06431 0.05934 -0.0300 0.0099 1.0000
17.000 1.5220 0.06807 0.06319 -0.0304 0.0097 1.0000
17.250 1.5163 0.07210 0.06731 -0.0309 0.0096 1.0000
17.500 1.5095 0.07637 0.07168 -0.0317 0.0095 1.0000
17.750 1.5018 0.08088 0.07629 -0.0327 0.0094 1.0000
18.000 1.4921 0.08581 0.08132 -0.0339 0.0093 1.0000
18.250 1.4815 0.09093 0.08653 -0.0353 0.0092 1.0000
18.500 1.4693 0.09640 0.09211 -0.0370 0.0091 1.0000
18.750 1.4561 0.10211 0.09792 -0.0388 0.0090 1.0000
19.000 1.4420 0.10803 0.10395 -0.0409 0.0090 1.0000
19.250 1.4275 0.11403 0.11005 -0.0431 0.0089 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 693 AIRFOIL (goe693-il)