GOE 116 (MVA MK.3) AIRFOIL (goe116-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 116 (MVA MK.3) AIRFOIL (goe116-il) Reynolds number: 1,000,000 Max Cl/Cd: 91.89 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe116-il-1000000-n5.txt Download as CSV file: xf-goe116-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 116 (MVA MK.3) 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4115 0.08618 0.08395 -0.0315 0.7709 0.0025
-8.750 -0.4131 0.08100 0.07877 -0.0344 0.7630 0.0025
-8.500 -0.4172 0.07516 0.07292 -0.0379 0.7561 0.0024
-8.250 -0.4271 0.06752 0.06529 -0.0434 0.7491 0.0023
-8.000 -0.4311 0.05490 0.05259 -0.0591 0.7430 0.0022
-7.750 -0.4176 0.03586 0.03308 -0.0825 0.7377 0.0022
-7.500 -0.4000 0.02924 0.02594 -0.0876 0.7317 0.0021
-7.250 -0.3800 0.02483 0.02105 -0.0896 0.7261 0.0021
-7.000 -0.3577 0.02146 0.01725 -0.0904 0.7200 0.0021
-6.750 -0.3337 0.01873 0.01410 -0.0907 0.7145 0.0021
-6.500 -0.3084 0.01648 0.01152 -0.0908 0.7092 0.0022
-6.250 -0.2823 0.01475 0.00951 -0.0909 0.7040 0.0025
-6.000 -0.2553 0.01364 0.00818 -0.0910 0.6993 0.0027
-5.750 -0.2277 0.01289 0.00728 -0.0911 0.6941 0.0029
-5.500 -0.2002 0.01196 0.00617 -0.0913 0.6889 0.0030
-5.250 -0.1723 0.01141 0.00550 -0.0916 0.6843 0.0033
-5.000 -0.1441 0.01108 0.00512 -0.0918 0.6796 0.0036
-4.750 -0.1159 0.01074 0.00469 -0.0921 0.6749 0.0040
-4.500 -0.0875 0.01028 0.00412 -0.0923 0.6706 0.0044
-4.250 -0.0589 0.00987 0.00360 -0.0925 0.6661 0.0047
-4.000 -0.0304 0.00945 0.00305 -0.0927 0.6617 0.0052
-3.750 -0.0019 0.00916 0.00268 -0.0929 0.6576 0.0059
-3.500 0.0268 0.00895 0.00243 -0.0931 0.6534 0.0070
-3.250 0.0554 0.00881 0.00224 -0.0933 0.6490 0.0079
-3.000 0.0840 0.00856 0.00187 -0.0935 0.6449 0.0103
-2.750 0.1128 0.00842 0.00166 -0.0937 0.6410 0.0132
-2.500 0.1415 0.00831 0.00150 -0.0939 0.6365 0.0211
-2.250 0.1694 0.00774 0.00132 -0.0944 0.6323 0.1242
-2.000 0.1981 0.00778 0.00133 -0.0946 0.6285 0.1316
-1.750 0.2269 0.00776 0.00131 -0.0948 0.6247 0.1361
-1.500 0.2556 0.00778 0.00135 -0.0951 0.6208 0.1412
-1.250 0.2843 0.00777 0.00130 -0.0953 0.6171 0.1424
-1.000 0.3130 0.00774 0.00126 -0.0955 0.6139 0.1432
-0.750 0.3417 0.00771 0.00121 -0.0958 0.6103 0.1439
-0.500 0.3704 0.00769 0.00118 -0.0960 0.6065 0.1446
-0.250 0.3991 0.00769 0.00116 -0.0963 0.6031 0.1452
0.000 0.4277 0.00768 0.00115 -0.0965 0.5999 0.1457
0.250 0.4564 0.00767 0.00115 -0.0967 0.5951 0.1463
0.500 0.4849 0.00766 0.00113 -0.0970 0.5894 0.1477
0.750 0.5135 0.00765 0.00113 -0.0972 0.5843 0.1492
1.000 0.5421 0.00763 0.00114 -0.0975 0.5793 0.1506
1.250 0.5704 0.00766 0.00116 -0.0977 0.5725 0.1520
1.500 0.5989 0.00766 0.00119 -0.0979 0.5639 0.1534
1.750 0.6272 0.00769 0.00123 -0.0981 0.5562 0.1548
2.000 0.6557 0.00771 0.00128 -0.0983 0.5495 0.1563
2.250 0.6839 0.00775 0.00134 -0.0985 0.5424 0.1578
2.500 0.7117 0.00786 0.00140 -0.0986 0.5202 0.1592
2.750 0.7388 0.00804 0.00150 -0.0986 0.4903 0.1604
3.000 0.7643 0.00845 0.00170 -0.0984 0.4339 0.1627
3.250 0.7797 0.01031 0.00255 -0.0971 0.2122 0.1652
3.500 0.7994 0.01150 0.00317 -0.0962 0.0757 0.1679
3.750 0.8230 0.01211 0.00360 -0.0957 0.0148 0.1711
4.000 0.8492 0.01236 0.00391 -0.0955 0.0091 0.1745
4.250 0.8756 0.01257 0.00418 -0.0954 0.0078 0.1811
4.750 0.9263 0.01319 0.00495 -0.0948 0.0053 0.2078
5.000 0.9523 0.01313 0.00534 -0.0949 0.0047 0.4176
5.500 0.9972 0.01270 0.00626 -0.0934 0.0037 1.0000
5.750 1.0194 0.01336 0.00698 -0.0925 0.0033 1.0000
6.000 1.0421 0.01393 0.00760 -0.0918 0.0031 1.0000
6.250 1.0637 0.01459 0.00834 -0.0908 0.0029 1.0000
6.500 1.0843 0.01530 0.00912 -0.0897 0.0027 1.0000
6.750 1.1055 0.01592 0.00979 -0.0888 0.0024 1.0000
7.000 1.1271 0.01645 0.01035 -0.0880 0.0022 1.0000
7.250 1.1477 0.01705 0.01101 -0.0871 0.0021 1.0000
7.500 1.1660 0.01782 0.01184 -0.0857 0.0020 1.0000
7.750 1.1803 0.01893 0.01304 -0.0838 0.0019 1.0000
8.000 1.1912 0.02037 0.01460 -0.0812 0.0018 1.0000
8.250 1.2059 0.02145 0.01578 -0.0793 0.0017 1.0000
8.500 1.2185 0.02271 0.01715 -0.0770 0.0016 1.0000
8.750 1.2305 0.02444 0.01905 -0.0745 0.0015 1.0000
9.000 1.2546 0.02941 0.02438 -0.0733 0.0013 1.0000
9.250 1.2729 0.03461 0.02998 -0.0713 0.0012 1.0000
9.500 1.2819 0.03914 0.03486 -0.0683 0.0012 1.0000
9.750 1.2852 0.04338 0.03942 -0.0650 0.0012 1.0000
10.000 1.2832 0.04739 0.04373 -0.0613 0.0013 1.0000
10.250 1.2764 0.05104 0.04762 -0.0573 0.0013 1.0000
10.500 1.2631 0.05418 0.05097 -0.0528 0.0013 1.0000
10.750 1.2479 0.05748 0.05446 -0.0490 0.0013 1.0000
11.000 1.2316 0.06111 0.05827 -0.0461 0.0014 1.0000
11.250 1.2144 0.06497 0.06229 -0.0441 0.0014 1.0000
11.500 1.1958 0.06926 0.06674 -0.0428 0.0014 1.0000
11.750 1.1770 0.07387 0.07151 -0.0423 0.0014 1.0000
12.000 1.1570 0.07897 0.07675 -0.0425 0.0014 1.0000
12.250 1.1372 0.08441 0.08233 -0.0435 0.0014 1.0000
12.500 1.1164 0.09038 0.08843 -0.0453 0.0014 1.0000
12.750 1.0950 0.09699 0.09516 -0.0479 0.0015 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 116 (MVA MK.3) AIRFOIL (goe116-il)