GOE 499 AIRFOIL (goe499-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 499 AIRFOIL (goe499-il) Reynolds number: 200,000 Max Cl/Cd: 102.16 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe499-il-200000.txt Download as CSV file: xf-goe499-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 499 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3149 0.10306 0.09980 -0.0350 0.9896 0.0379
-7.500 -0.3034 0.10027 0.09702 -0.0402 0.9854 0.0400
-7.250 -0.3027 0.09909 0.09586 -0.0453 0.9783 0.0409
-7.000 -0.2824 0.09548 0.09225 -0.0539 0.9727 0.0413
-6.750 -0.2702 0.08894 0.08573 -0.0568 0.9707 0.0423
-6.500 -0.2589 0.08654 0.08331 -0.0528 0.9690 0.0445
-6.250 -0.2522 0.08427 0.08104 -0.0531 0.9634 0.0465
-6.000 -0.2319 0.08072 0.07747 -0.0584 0.9593 0.0497
-5.750 -0.1931 0.07612 0.07281 -0.0784 0.9517 0.0540
-5.500 -0.1853 0.07106 0.06776 -0.0789 0.9478 0.0553
-5.250 -0.1713 0.06895 0.06563 -0.0766 0.9455 0.0571
-5.000 -0.1414 0.06593 0.06256 -0.0816 0.9435 0.0630
-4.750 -0.1078 0.06010 0.05664 -0.0948 0.9372 0.0688
-4.500 -0.0458 0.05524 0.05150 -0.1117 0.9338 0.0801
-4.250 -0.0287 0.05144 0.04776 -0.1115 0.9314 0.0820
-4.000 -0.0010 0.04975 0.04605 -0.1121 0.9294 0.0860
-3.750 0.0429 0.04509 0.04119 -0.1218 0.9238 0.0970
-3.250 0.1993 0.02592 0.01991 -0.1537 0.9222 0.0677
-3.000 0.2421 0.02330 0.01694 -0.1573 0.9195 0.0654
-2.750 0.2875 0.02108 0.01427 -0.1610 0.9175 0.0632
-2.500 0.3331 0.01956 0.01241 -0.1643 0.9160 0.0628
-2.250 0.3651 0.01874 0.01143 -0.1650 0.9114 0.0634
-2.000 0.3952 0.01813 0.01073 -0.1654 0.9068 0.0644
-1.750 0.4327 0.01759 0.01010 -0.1672 0.9045 0.0669
-1.500 0.4733 0.01683 0.00938 -0.1697 0.9029 0.0707
-1.250 0.5144 0.01631 0.00885 -0.1722 0.9015 0.0752
-1.000 0.5358 0.01626 0.00881 -0.1709 0.8950 0.0818
-0.750 0.5708 0.01571 0.00855 -0.1723 0.8918 0.1493
-0.500 0.6081 0.01535 0.00839 -0.1740 0.8894 0.2393
-0.250 0.6466 0.01499 0.00814 -0.1759 0.8874 0.2866
0.000 0.6689 0.01498 0.00828 -0.1749 0.8804 0.3289
0.250 0.7024 0.01455 0.00818 -0.1759 0.8766 0.4158
0.500 0.7376 0.01401 0.00803 -0.1770 0.8740 0.5649
0.750 0.7615 0.01321 0.00796 -0.1757 0.8685 1.0000
1.000 0.7919 0.01312 0.00779 -0.1759 0.8629 1.0000
1.250 0.8280 0.01285 0.00745 -0.1770 0.8596 1.0000
1.500 0.8528 0.01290 0.00748 -0.1762 0.8515 1.0000
1.750 0.8868 0.01260 0.00717 -0.1768 0.8464 1.0000
2.000 0.9137 0.01252 0.00708 -0.1762 0.8378 1.0000
2.250 0.9472 0.01218 0.00674 -0.1766 0.8314 1.0000
2.500 0.9727 0.01213 0.00672 -0.1757 0.8212 1.0000
2.750 1.0020 0.01195 0.00659 -0.1755 0.8124 1.0000
3.000 1.0292 0.01159 0.00622 -0.1744 0.7961 1.0000
3.250 1.0570 0.01112 0.00571 -0.1733 0.7742 1.0000
3.500 1.0789 0.01085 0.00539 -0.1711 0.7375 1.0000
3.750 1.1023 0.01079 0.00526 -0.1695 0.6981 1.0000
4.000 1.1225 0.01102 0.00509 -0.1671 0.6197 1.0000
4.250 1.1323 0.01212 0.00543 -0.1631 0.5113 1.0000
4.500 1.1381 0.01358 0.00615 -0.1589 0.3953 1.0000
4.750 1.1276 0.01689 0.00773 -0.1530 0.1076 1.0000
5.000 1.1404 0.01836 0.00883 -0.1504 0.0435 1.0000
5.250 1.1580 0.01933 0.00992 -0.1484 0.0376 1.0000
5.500 1.1765 0.02014 0.01085 -0.1466 0.0343 1.0000
5.750 1.1924 0.02113 0.01195 -0.1444 0.0323 1.0000
6.000 1.2060 0.02228 0.01318 -0.1419 0.0310 1.0000
6.250 1.2176 0.02355 0.01452 -0.1390 0.0301 1.0000
6.500 1.2290 0.02497 0.01598 -0.1361 0.0296 1.0000
6.750 1.2436 0.02653 0.01760 -0.1338 0.0293 1.0000
7.000 1.2676 0.02900 0.02004 -0.1333 0.0278 1.0000
7.250 1.3100 0.03248 0.02355 -0.1355 0.0272 1.0000
7.500 1.3386 0.03401 0.02519 -0.1351 0.0277 1.0000
7.750 1.3675 0.03538 0.02689 -0.1340 0.0302 1.0000
12.500 1.2673 0.11986 0.11684 -0.0987 0.0458 1.0000
12.750 1.2451 0.12819 0.12527 -0.1036 0.0464 1.0000
13.000 0.9902 0.15711 0.15471 -0.1186 0.0681 1.0000
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Polar data table (+)
Polar graphs
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