GOE 447 AIRFOIL (goe447-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 447 AIRFOIL (goe447-il) Reynolds number: 1,000,000 Max Cl/Cd: 161.26 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe447-il-1000000.txt Download as CSV file: xf-goe447-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 447 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
------ -------- --------- --------- -------- -------- --------
-9.250 0.0391 0.08752 0.08471 -0.1124 0.7674 0.0256
-9.000 0.0458 0.08446 0.08161 -0.1144 0.7633 0.0256
-8.750 0.0525 0.08102 0.07817 -0.1153 0.7598 0.0259
-8.500 0.0650 0.07901 0.07617 -0.1160 0.7563 0.0260
-8.250 0.0771 0.07703 0.07418 -0.1169 0.7524 0.0262
-8.000 0.0886 0.07503 0.07216 -0.1180 0.7485 0.0264
-7.750 0.0998 0.07296 0.07006 -0.1193 0.7442 0.0267
-7.500 0.1107 0.07077 0.06789 -0.1209 0.7403 0.0272
-7.250 0.1197 0.06851 0.06563 -0.1227 0.7358 0.0280
-6.750 0.1461 0.05779 0.05482 -0.1388 0.7273 0.0301
-6.500 0.1445 0.01692 0.01241 -0.1791 0.7252 0.0286
-6.250 0.1701 0.01540 0.01061 -0.1796 0.7213 0.0288
-6.000 0.1967 0.01446 0.00946 -0.1798 0.7174 0.0291
-5.750 0.2239 0.01393 0.00875 -0.1800 0.7134 0.0295
-5.500 0.2518 0.01350 0.00819 -0.1801 0.7100 0.0297
-5.250 0.2787 0.01248 0.00699 -0.1803 0.7066 0.0300
-5.000 0.3059 0.01167 0.00603 -0.1805 0.7029 0.0304
-4.750 0.3335 0.01120 0.00547 -0.1806 0.6991 0.0308
-4.500 0.3613 0.01088 0.00507 -0.1807 0.6952 0.0312
-4.250 0.3897 0.01055 0.00471 -0.1809 0.6924 0.0316
-4.000 0.4180 0.01027 0.00439 -0.1810 0.6889 0.0320
-3.750 0.4462 0.01003 0.00409 -0.1811 0.6851 0.0325
-3.500 0.4742 0.00983 0.00381 -0.1812 0.6810 0.0330
-3.250 0.5025 0.00964 0.00358 -0.1813 0.6775 0.0335
-3.000 0.5311 0.00946 0.00336 -0.1814 0.6742 0.0341
-2.750 0.5594 0.00925 0.00312 -0.1816 0.6707 0.0348
-2.500 0.5876 0.00905 0.00289 -0.1817 0.6675 0.0362
-2.250 0.6157 0.00895 0.00276 -0.1818 0.6642 0.0373
-2.000 0.6442 0.00885 0.00264 -0.1819 0.6612 0.0385
-1.750 0.6728 0.00873 0.00252 -0.1821 0.6580 0.0399
-1.500 0.7011 0.00861 0.00242 -0.1822 0.6545 0.0432
-1.250 0.7292 0.00855 0.00238 -0.1823 0.6511 0.0491
-1.000 0.7573 0.00861 0.00245 -0.1823 0.6476 0.0596
-0.750 0.7858 0.00868 0.00256 -0.1825 0.6447 0.0659
-0.500 0.8142 0.00870 0.00259 -0.1826 0.6416 0.0697
-0.250 0.8424 0.00872 0.00264 -0.1828 0.6380 0.0731
0.000 0.8700 0.00877 0.00267 -0.1828 0.6338 0.0758
0.250 0.8975 0.00884 0.00271 -0.1828 0.6295 0.0775
0.500 0.9255 0.00873 0.00261 -0.1829 0.6255 0.0802
0.750 0.9532 0.00869 0.00259 -0.1829 0.6212 0.0827
1.000 0.9802 0.00873 0.00261 -0.1829 0.6164 0.0854
1.250 1.0076 0.00874 0.00263 -0.1829 0.6115 0.0877
1.500 1.0350 0.00876 0.00265 -0.1829 0.6060 0.0893
1.750 1.0618 0.00874 0.00261 -0.1827 0.6009 0.0922
2.000 1.0888 0.00875 0.00262 -0.1827 0.5963 0.0947
2.250 1.1163 0.00874 0.00264 -0.1827 0.5913 0.0970
2.500 1.1428 0.00879 0.00268 -0.1826 0.5856 0.0993
2.750 1.1690 0.00885 0.00273 -0.1824 0.5800 0.1018
3.000 1.1956 0.00888 0.00278 -0.1822 0.5727 0.1078
3.250 1.2207 0.00895 0.00286 -0.1818 0.5646 0.1273
3.500 1.2468 0.00881 0.00306 -0.1818 0.5559 0.3011
3.750 1.2709 0.00820 0.00338 -0.1817 0.5472 0.7833
4.000 1.2917 0.00801 0.00351 -0.1802 0.5367 1.0000
4.250 1.3155 0.00820 0.00364 -0.1795 0.5242 1.0000
4.500 1.3378 0.00844 0.00381 -0.1786 0.5095 1.0000
4.750 1.3591 0.00872 0.00400 -0.1775 0.4946 1.0000
5.000 1.3792 0.00903 0.00424 -0.1762 0.4793 1.0000
5.250 1.3984 0.00937 0.00449 -0.1747 0.4648 1.0000
5.500 1.4156 0.00971 0.00476 -0.1727 0.4525 1.0000
5.750 1.4343 0.01002 0.00502 -0.1712 0.4426 1.0000
6.000 1.4534 0.01033 0.00530 -0.1697 0.4343 1.0000
6.250 1.4723 0.01065 0.00559 -0.1682 0.4262 1.0000
6.500 1.4917 0.01097 0.00589 -0.1669 0.4196 1.0000
6.750 1.5124 0.01124 0.00616 -0.1658 0.4139 1.0000
7.250 1.5498 0.01196 0.00685 -0.1630 0.3992 1.0000
7.500 1.5672 0.01238 0.00724 -0.1615 0.3920 1.0000
7.750 1.5863 0.01274 0.00761 -0.1602 0.3854 1.0000
8.000 1.6033 0.01320 0.00805 -0.1587 0.3764 1.0000
8.250 1.6210 0.01365 0.00849 -0.1573 0.3683 1.0000
8.500 1.6359 0.01424 0.00904 -0.1555 0.3579 1.0000
8.750 1.6521 0.01479 0.00958 -0.1540 0.3462 1.0000
9.000 1.6642 0.01556 0.01030 -0.1519 0.3305 1.0000
9.250 1.6705 0.01668 0.01130 -0.1490 0.3072 1.0000
9.500 1.6682 0.01838 0.01280 -0.1451 0.2723 1.0000
9.750 1.6631 0.02036 0.01459 -0.1411 0.2399 1.0000
10.000 1.6623 0.02217 0.01629 -0.1379 0.2173 1.0000
10.250 1.6663 0.02374 0.01779 -0.1354 0.2018 1.0000
10.500 1.6714 0.02524 0.01924 -0.1331 0.1899 1.0000
10.750 1.6794 0.02659 0.02058 -0.1313 0.1809 1.0000
11.000 1.6865 0.02805 0.02201 -0.1294 0.1724 1.0000
11.250 1.6933 0.02953 0.02348 -0.1275 0.1643 1.0000
11.500 1.7017 0.03094 0.02490 -0.1259 0.1573 1.0000
11.750 1.7066 0.03265 0.02658 -0.1240 0.1483 1.0000
12.000 1.7118 0.03436 0.02826 -0.1222 0.1364 1.0000
12.250 1.6996 0.03755 0.03125 -0.1190 0.1029 1.0000
12.500 1.6798 0.04159 0.03513 -0.1154 0.0755 1.0000
12.750 1.6687 0.04499 0.03847 -0.1128 0.0567 1.0000
13.000 1.6498 0.04925 0.04258 -0.1098 0.0308 1.0000
13.250 1.6475 0.05204 0.04537 -0.1082 0.0249 1.0000
13.500 1.6502 0.05439 0.04775 -0.1069 0.0226 1.0000
13.750 1.6511 0.05693 0.05034 -0.1056 0.0208 1.0000
14.000 1.6560 0.05908 0.05255 -0.1046 0.0198 1.0000
14.250 1.6581 0.06158 0.05510 -0.1036 0.0189 1.0000
14.500 1.6593 0.06420 0.05777 -0.1025 0.0180 1.0000
14.750 1.6590 0.06703 0.06067 -0.1015 0.0173 1.0000
15.000 1.6622 0.06949 0.06319 -0.1007 0.0168 1.0000
15.250 1.6644 0.07206 0.06583 -0.0999 0.0164 1.0000
15.500 1.6654 0.07483 0.06866 -0.0992 0.0159 1.0000
15.750 1.6660 0.07765 0.07154 -0.0985 0.0155 1.0000
16.000 1.6638 0.08088 0.07484 -0.0979 0.0150 1.0000
16.250 1.6592 0.08443 0.07847 -0.0972 0.0146 1.0000
16.500 1.6550 0.08794 0.08207 -0.0967 0.0142 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 447 AIRFOIL (goe447-il)