GOE 115 (MVA MK.2) AIRFOIL (goe115-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 115 (MVA MK.2) AIRFOIL (goe115-il) Reynolds number: 1,000,000 Max Cl/Cd: 131.98 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe115-il-1000000.txt Download as CSV file: xf-goe115-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3251 0.08013 0.07866 -0.0418 1.0000 0.0127
-9.750 -0.3259 0.07666 0.07520 -0.0424 1.0000 0.0129
-9.250 -0.4497 0.04211 0.04071 -0.0530 1.0000 0.0092
-9.000 -0.4323 0.04331 0.04195 -0.0518 1.0000 0.0095
-8.750 -0.5171 0.03683 0.03488 -0.0811 0.9874 0.0085
-8.500 -0.5163 0.02199 0.01877 -0.0910 0.9731 0.0094
-8.250 -0.4823 0.02148 0.01815 -0.0932 0.9672 0.0097
-8.000 -0.4498 0.02048 0.01698 -0.0952 0.9578 0.0101
-7.750 -0.4201 0.01875 0.01500 -0.0966 0.9446 0.0107
-7.500 -0.3941 0.01681 0.01275 -0.0970 0.9279 0.0115
-7.250 -0.3699 0.01561 0.01130 -0.0967 0.9102 0.0121
-7.000 -0.3486 0.01403 0.00944 -0.0959 0.8930 0.0130
-6.750 -0.3225 0.01407 0.00941 -0.0957 0.8779 0.0135
-6.500 -0.2960 0.01425 0.00951 -0.0955 0.8638 0.0142
-6.250 -0.2709 0.01385 0.00899 -0.0951 0.8505 0.0151
-6.000 -0.2457 0.01338 0.00837 -0.0946 0.8380 0.0161
-5.750 -0.2194 0.01326 0.00811 -0.0942 0.8261 0.0168
-5.500 -0.1966 0.01193 0.00666 -0.0935 0.8144 0.0181
-5.250 -0.1707 0.01166 0.00634 -0.0933 0.8028 0.0191
-5.000 -0.1445 0.01148 0.00609 -0.0930 0.7912 0.0203
-4.750 -0.1185 0.01115 0.00567 -0.0927 0.7792 0.0215
-4.500 -0.0922 0.01090 0.00533 -0.0923 0.7666 0.0226
-4.250 -0.0651 0.01093 0.00526 -0.0921 0.7530 0.0234
-4.000 -0.0420 0.00983 0.00406 -0.0914 0.7389 0.0252
-3.750 -0.0166 0.00939 0.00355 -0.0909 0.7241 0.0264
-3.500 0.0093 0.00908 0.00315 -0.0906 0.7094 0.0276
-3.250 0.0355 0.00884 0.00281 -0.0902 0.6952 0.0287
-3.000 0.0620 0.00865 0.00253 -0.0900 0.6819 0.0301
-2.750 0.0887 0.00848 0.00227 -0.0897 0.6699 0.0308
-2.500 0.1156 0.00833 0.00203 -0.0895 0.6593 0.0313
-2.250 0.1426 0.00820 0.00183 -0.0893 0.6501 0.0316
-2.000 0.1697 0.00810 0.00165 -0.0891 0.6414 0.0320
-1.750 0.1970 0.00800 0.00151 -0.0889 0.6338 0.0326
-1.500 0.2241 0.00791 0.00132 -0.0887 0.6266 0.0341
-1.250 0.2517 0.00784 0.00120 -0.0886 0.6197 0.0363
-1.000 0.2789 0.00782 0.00112 -0.0885 0.6129 0.0389
-0.750 0.3065 0.00777 0.00107 -0.0884 0.6068 0.0424
-0.500 0.3325 0.00739 0.00096 -0.0881 0.6010 0.1491
-0.250 0.3584 0.00712 0.00096 -0.0879 0.5960 0.2409
0.000 0.3860 0.00708 0.00101 -0.0879 0.5913 0.2736
0.250 0.4135 0.00710 0.00106 -0.0878 0.5864 0.2905
0.500 0.4409 0.00714 0.00110 -0.0877 0.5817 0.3004
0.750 0.4687 0.00712 0.00113 -0.0877 0.5775 0.3094
1.000 0.4964 0.00715 0.00116 -0.0877 0.5732 0.3152
1.250 0.5238 0.00719 0.00120 -0.0876 0.5693 0.3202
1.500 0.5514 0.00720 0.00125 -0.0875 0.5654 0.3263
1.750 0.5791 0.00721 0.00129 -0.0875 0.5609 0.3330
2.000 0.6064 0.00722 0.00134 -0.0874 0.5566 0.3438
2.250 0.6336 0.00725 0.00141 -0.0873 0.5527 0.3538
2.500 0.6613 0.00724 0.00148 -0.0873 0.5494 0.3635
2.750 0.6884 0.00724 0.00153 -0.0871 0.5432 0.3785
3.000 0.7151 0.00719 0.00161 -0.0869 0.5345 0.4117
3.250 0.7396 0.00693 0.00171 -0.0864 0.5225 0.5642
3.500 0.7865 0.00610 0.00187 -0.0909 0.5074 1.0000
3.750 0.8125 0.00621 0.00195 -0.0905 0.4951 1.0000
4.000 0.8381 0.00635 0.00204 -0.0901 0.4723 1.0000
4.250 0.8605 0.00674 0.00219 -0.0891 0.4133 1.0000
4.500 0.8745 0.00790 0.00274 -0.0868 0.2907 1.0000
4.750 0.8778 0.01006 0.00382 -0.0830 0.0697 1.0000
5.000 0.8982 0.01070 0.00428 -0.0817 0.0300 1.0000
5.250 0.9211 0.01110 0.00466 -0.0808 0.0214 1.0000
5.500 0.9438 0.01151 0.00506 -0.0799 0.0175 1.0000
5.750 0.9663 0.01194 0.00554 -0.0789 0.0158 1.0000
6.000 0.9893 0.01230 0.00593 -0.0781 0.0147 1.0000
6.250 1.0116 0.01271 0.00637 -0.0771 0.0136 1.0000
6.500 1.0325 0.01322 0.00691 -0.0760 0.0125 1.0000
6.750 1.0476 0.01417 0.00797 -0.0738 0.0113 1.0000
7.000 1.0694 0.01455 0.00838 -0.0728 0.0109 1.0000
7.250 1.0888 0.01510 0.00897 -0.0714 0.0104 1.0000
7.500 1.1070 0.01570 0.00964 -0.0698 0.0099 1.0000
7.750 1.1242 0.01634 0.01033 -0.0681 0.0094 1.0000
8.000 1.1404 0.01703 0.01106 -0.0663 0.0091 1.0000
8.250 1.1564 0.01768 0.01173 -0.0645 0.0086 1.0000
8.500 1.1665 0.01862 0.01274 -0.0617 0.0084 1.0000
8.750 1.1692 0.01997 0.01416 -0.0576 0.0080 1.0000
9.000 1.1723 0.02195 0.01626 -0.0539 0.0077 1.0000
9.250 1.1874 0.02249 0.01687 -0.0521 0.0075 1.0000
9.500 1.2020 0.02312 0.01758 -0.0503 0.0071 1.0000
9.750 1.2147 0.02422 0.01876 -0.0483 0.0069 1.0000
10.000 1.2276 0.02546 0.02009 -0.0465 0.0067 1.0000
10.250 1.2406 0.02682 0.02156 -0.0447 0.0065 1.0000
10.500 1.2539 0.02843 0.02330 -0.0431 0.0064 1.0000
10.750 1.2669 0.03036 0.02539 -0.0415 0.0063 1.0000
11.000 1.2760 0.03142 0.02652 -0.0397 0.0061 1.0000
11.250 1.2856 0.03401 0.02935 -0.0378 0.0060 1.0000
11.500 1.2912 0.03427 0.02959 -0.0361 0.0058 1.0000
11.750 1.2949 0.03790 0.03353 -0.0339 0.0058 1.0000
12.000 1.2985 0.03871 0.03435 -0.0324 0.0056 1.0000
12.250 1.3014 0.04013 0.03585 -0.0311 0.0055 1.0000
12.500 1.2919 0.04501 0.04110 -0.0287 0.0056 1.0000
12.750 1.2911 0.04690 0.04306 -0.0277 0.0055 1.0000
13.000 1.2815 0.05060 0.04696 -0.0265 0.0054 1.0000
13.250 1.2386 0.05882 0.05556 -0.0256 0.0052 1.0000
13.500 1.2509 0.05952 0.05627 -0.0255 0.0053 1.0000
13.750 1.2147 0.06742 0.06448 -0.0263 0.0052 1.0000
14.000 1.2099 0.07094 0.06811 -0.0270 0.0052 1.0000
14.250 1.1895 0.07713 0.07449 -0.0288 0.0052 1.0000
14.500 1.1602 0.08528 0.08285 -0.0318 0.0052 1.0000
14.750 1.1491 0.09087 0.08857 -0.0342 0.0053 1.0000
15.000 1.1181 0.10065 0.09853 -0.0392 0.0052 1.0000
15.250 1.1009 0.10854 0.10657 -0.0435 0.0053 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 115 (MVA MK.2) AIRFOIL (goe115-il)