GOE 115 (MVA MK.2) AIRFOIL (goe115-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 115 (MVA MK.2) AIRFOIL (goe115-il) Reynolds number: 100,000 Max Cl/Cd: 57.1 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe115-il-100000.txt Download as CSV file: xf-goe115-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 115 (MVA MK.2) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3996 0.09885 0.09431 -0.0394 1.0000 0.1340
-8.250 -0.3558 0.09299 0.08829 -0.0343 1.0000 0.1431
-8.000 -0.3809 0.09207 0.08759 -0.0364 1.0000 0.1473
-7.750 -0.3577 0.08775 0.08322 -0.0336 1.0000 0.1547
-7.500 -0.3775 0.08664 0.08228 -0.0334 1.0000 0.1607
-7.250 -0.4115 0.08583 0.08170 -0.0365 1.0000 0.1622
-7.000 -0.3811 0.08162 0.07743 -0.0294 1.0000 0.1686
-6.750 -0.4000 0.08046 0.07642 -0.0282 1.0000 0.1741
-6.500 -0.4313 0.07911 0.07522 -0.0293 1.0000 0.1775
-6.250 -0.4260 0.07696 0.07312 -0.0233 1.0000 0.1805
-6.000 -0.4352 0.07554 0.07177 -0.0208 1.0000 0.1854
-5.750 -0.4285 0.07187 0.06800 -0.0353 0.9942 0.2041
-5.500 -0.3905 0.03994 0.03447 -0.0661 0.9856 0.0961
-5.250 -0.3501 0.03320 0.02669 -0.0704 0.9785 0.0786
-5.000 -0.3089 0.02918 0.02194 -0.0739 0.9716 0.0749
-4.750 -0.2697 0.02677 0.01896 -0.0765 0.9630 0.0763
-4.500 -0.2254 0.02454 0.01625 -0.0794 0.9567 0.0757
-4.250 -0.1875 0.02302 0.01448 -0.0811 0.9473 0.0777
-4.000 -0.1414 0.02169 0.01294 -0.0844 0.9416 0.0819
-3.750 -0.1070 0.02047 0.01177 -0.0857 0.9314 0.0850
-3.500 -0.0658 0.01960 0.01081 -0.0882 0.9240 0.0913
-3.250 -0.0258 0.01876 0.00996 -0.0907 0.9161 0.1046
-3.000 0.0108 0.01773 0.00935 -0.0925 0.9077 0.1742
-2.750 0.0522 0.01795 0.00970 -0.0947 0.9002 0.3081
-2.500 0.0837 0.01824 0.00982 -0.0952 0.8898 0.3461
-2.250 0.1226 0.01818 0.00975 -0.0970 0.8841 0.3821
-2.000 0.1495 0.01818 0.00966 -0.0969 0.8732 0.3999
-1.750 0.1794 0.01811 0.00954 -0.0972 0.8647 0.4197
-1.500 0.2120 0.01792 0.00928 -0.0980 0.8573 0.4343
-1.250 0.2388 0.01789 0.00922 -0.0978 0.8481 0.4484
-1.000 0.2699 0.01771 0.00912 -0.0982 0.8418 0.4789
-0.750 0.2928 0.01778 0.00931 -0.0974 0.8326 0.5079
-0.500 0.3238 0.01760 0.00916 -0.0978 0.8272 0.5305
-0.250 0.3474 0.01771 0.00933 -0.0973 0.8184 0.5502
0.000 0.3792 0.01751 0.00921 -0.0978 0.8130 0.5766
0.250 0.4031 0.01754 0.00946 -0.0974 0.8050 0.6129
0.750 0.4867 0.01678 0.00941 -0.1027 0.7947 1.0000
1.000 0.5117 0.01718 0.00968 -0.1025 0.7876 1.0000
1.250 0.5414 0.01738 0.00977 -0.1027 0.7832 1.0000
1.500 0.5617 0.01800 0.01036 -0.1019 0.7751 1.0000
1.750 0.5908 0.01813 0.01042 -0.1018 0.7696 1.0000
2.000 0.6124 0.01857 0.01085 -0.1009 0.7604 1.0000
2.250 0.6424 0.01852 0.01076 -0.1005 0.7537 1.0000
2.500 0.6642 0.01893 0.01119 -0.0996 0.7442 1.0000
2.750 0.6918 0.01913 0.01140 -0.0993 0.7387 1.0000
3.000 0.7135 0.01968 0.01201 -0.0985 0.7307 1.0000
3.250 0.7431 0.01979 0.01217 -0.0984 0.7258 1.0000
3.500 0.7624 0.02048 0.01297 -0.0974 0.7170 1.0000
3.750 0.7921 0.02057 0.01312 -0.0972 0.7119 1.0000
4.000 0.8115 0.02127 0.01394 -0.0962 0.7028 1.0000
4.250 0.8420 0.02130 0.01408 -0.0960 0.6974 1.0000
4.500 0.8621 0.02187 0.01482 -0.0949 0.6871 1.0000
4.750 0.8950 0.02072 0.01369 -0.0932 0.6699 1.0000
5.000 0.9240 0.01925 0.01215 -0.0904 0.6413 1.0000
5.250 0.9483 0.01821 0.01110 -0.0875 0.6085 1.0000
5.500 0.9669 0.01726 0.01014 -0.0837 0.5605 1.0000
5.750 0.9667 0.01693 0.00905 -0.0765 0.3746 1.0000
6.000 0.9467 0.02062 0.01075 -0.0698 0.1106 1.0000
6.500 0.9623 0.02394 0.01390 -0.0635 0.0729 1.0000
6.750 0.9687 0.02549 0.01551 -0.0603 0.0673 1.0000
7.000 0.9780 0.02684 0.01696 -0.0575 0.0620 1.0000
7.250 0.9878 0.02841 0.01849 -0.0548 0.0593 1.0000
7.500 1.0173 0.03121 0.02100 -0.0547 0.0565 1.0000
7.750 1.0479 0.03288 0.02281 -0.0545 0.0546 1.0000
8.000 1.0778 0.03477 0.02486 -0.0542 0.0517 1.0000
8.250 1.1132 0.03749 0.02780 -0.0546 0.0515 1.0000
8.500 1.1441 0.04066 0.03126 -0.0543 0.0528 1.0000
8.750 1.1725 0.04476 0.03563 -0.0540 0.0550 1.0000
9.000 1.2004 0.04766 0.03880 -0.0531 0.0582 1.0000
9.250 1.2085 0.05067 0.04268 -0.0491 0.0641 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 115 (MVA MK.2) AIRFOIL (goe115-il)