GOE 243 (MVA PR.3) AIRFOIL (goe243-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 243 (MVA PR.3) AIRFOIL (goe243-il) Reynolds number: 50,000 Max Cl/Cd: 17.93 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe243-il-50000-n5.txt Download as CSV file: xf-goe243-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 243 (MVA PR.3) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 0.1110 0.12003 0.11293 -0.0974 0.8714 0.1687
-8.250 0.0870 0.11018 0.10297 -0.1040 0.8572 0.1010
-7.750 0.0865 0.10284 0.09559 -0.1077 0.8442 0.0902
-7.500 0.0834 0.10200 0.09481 -0.1047 0.8344 0.0895
-7.250 0.0917 0.09959 0.09241 -0.1048 0.8280 0.0883
-7.000 0.1025 0.09610 0.08891 -0.1069 0.8237 0.0863
-6.750 0.0780 0.09566 0.08855 -0.1022 0.8110 0.0851
-6.500 0.0521 0.08980 0.08268 -0.1050 0.8036 0.0811
-6.250 0.0322 0.08941 0.08237 -0.1005 0.7915 0.0807
-6.000 0.0361 0.08594 0.07890 -0.1021 0.7850 0.0798
-5.750 0.0450 0.08012 0.07301 -0.1081 0.7809 0.0786
-5.500 0.0149 0.07821 0.07115 -0.1060 0.7666 0.0777
-5.250 0.0577 0.06153 0.05362 -0.1415 0.7617 0.0752
-5.000 0.0944 0.05753 0.04921 -0.1517 0.7539 0.0757
-4.750 0.1324 0.05488 0.04623 -0.1586 0.7465 0.0773
-4.500 0.1889 0.05178 0.04260 -0.1674 0.7426 0.0796
-4.250 0.2447 0.04925 0.03957 -0.1741 0.7398 0.0812
-4.000 0.2558 0.04924 0.03950 -0.1731 0.7288 0.0818
-3.750 0.2866 0.04820 0.03841 -0.1735 0.7232 0.0831
-3.500 0.3237 0.04695 0.03708 -0.1743 0.7197 0.0848
-3.250 0.3376 0.04716 0.03725 -0.1730 0.7099 0.0863
-3.000 0.3662 0.04667 0.03662 -0.1732 0.7031 0.0893
-2.750 0.4017 0.04566 0.03563 -0.1736 0.6994 0.0931
-2.500 0.4195 0.04590 0.03588 -0.1727 0.6904 0.0963
-2.250 0.4448 0.04581 0.03572 -0.1724 0.6827 0.0997
-2.000 0.4848 0.04482 0.03481 -0.1740 0.6791 0.1050
-1.750 0.5132 0.04482 0.03481 -0.1753 0.6710 0.1123
-1.500 0.5510 0.04448 0.03457 -0.1789 0.6627 0.1246
-1.250 0.6118 0.04294 0.03428 -0.1860 0.6593 0.2408
-0.750 0.6479 0.04540 0.03650 -0.1827 0.6417 0.3719
-0.500 0.6678 0.04594 0.03699 -0.1786 0.6376 0.3997
-0.250 0.6675 0.04811 0.03918 -0.1753 0.6251 0.4140
0.250 0.7110 0.04875 0.03967 -0.1691 0.6164 0.4562
0.750 0.7332 0.05094 0.04180 -0.1648 0.5985 0.4701
1.000 0.7886 0.04967 0.04022 -0.1698 0.5960 0.4769
1.500 0.8191 0.05183 0.04230 -0.1686 0.5783 0.4823
1.750 0.8611 0.05081 0.04112 -0.1701 0.5758 0.4863
2.250 0.8930 0.05367 0.04389 -0.1707 0.5582 0.4947
2.500 0.9386 0.05265 0.04268 -0.1730 0.5560 0.5016
3.000 0.9540 0.05650 0.04657 -0.1708 0.5382 0.5089
3.250 0.9968 0.05558 0.04550 -0.1725 0.5361 0.5164
4.750 1.0669 0.06691 0.05679 -0.1698 0.4941 0.5486
5.250 1.0847 0.07157 0.06143 -0.1693 0.4786 0.5618
5.500 1.1243 0.06969 0.05948 -0.1687 0.4768 0.5699
6.000 1.1244 0.07610 0.06595 -0.1673 0.4585 0.5827
6.250 1.1698 0.07323 0.06297 -0.1667 0.4568 0.5932
6.750 1.1717 0.07892 0.06873 -0.1649 0.4382 0.6054
7.250 1.1771 0.08428 0.07412 -0.1634 0.4198 0.6199
7.500 1.2107 0.08261 0.07239 -0.1624 0.4169 0.6302
8.000 1.1537 0.09719 0.08726 -0.1617 0.3895 0.6351
8.250 1.1693 0.09819 0.08827 -0.1610 0.3838 0.6434
8.500 1.1993 0.09715 0.08716 -0.1603 0.3810 0.6558
8.750 1.2332 0.09530 0.08529 -0.1592 0.3790 0.6674
9.250 1.2097 0.10538 0.09552 -0.1591 0.3568 0.6774
9.750 1.2279 0.10906 0.09926 -0.1582 0.3421 0.6948
10.000 1.2040 0.11623 0.10656 -0.1590 0.3271 0.6973
10.250 1.2296 0.11550 0.10580 -0.1581 0.3242 0.7094
10.500 1.2581 0.11422 0.10447 -0.1570 0.3222 0.7214
11.000 1.2516 0.12234 0.11271 -0.1578 0.3036 0.7349
11.500 1.2378 0.13242 0.12290 -0.1599 0.2847 0.7470
12.000 1.2444 0.13909 0.12963 -0.1612 0.2722 0.7633
12.250 1.2669 0.13869 0.12922 -0.1605 0.2706 0.7768
13.250 1.2143 0.16764 0.15848 -0.1718 0.2458 0.7968
13.500 1.2204 0.17098 0.16188 -0.1731 0.2434 0.8086
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