GOE 310 (MVA H.42) AIRFOIL (goe310-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 310 (MVA H.42) AIRFOIL (goe310-il) Reynolds number: 100,000 Max Cl/Cd: 59.59 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe310-il-100000.txt Download as CSV file: xf-goe310-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 310 (MVA H.42) 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.2520 0.10333 0.09869 -0.0339 1.0000 0.0898
-9.250 -0.2647 0.10234 0.09779 -0.0350 1.0000 0.0921
-9.000 -0.2828 0.10191 0.09746 -0.0360 1.0000 0.0927
-8.750 -0.2632 0.09534 0.09090 -0.0336 1.0000 0.0951
-8.500 -0.2566 0.09222 0.08780 -0.0319 1.0000 0.0984
-8.250 -0.2597 0.08995 0.08560 -0.0309 1.0000 0.1017
-8.000 -0.3347 0.09627 0.09160 -0.0309 1.0000 0.0965
-7.750 -0.3350 0.09403 0.08943 -0.0292 1.0000 0.0997
-7.500 -0.3448 0.09254 0.08804 -0.0275 1.0000 0.1021
-7.250 -0.3624 0.09167 0.08728 -0.0251 1.0000 0.1039
-7.000 -0.3853 0.09165 0.08736 -0.0256 1.0000 0.1059
-6.750 -0.4040 0.09214 0.08786 -0.0293 1.0000 0.1067
-6.500 -0.3916 0.08572 0.08156 -0.0216 1.0000 0.1091
-6.250 -0.3905 0.08345 0.07933 -0.0186 1.0000 0.1129
-6.000 -0.3966 0.08187 0.07779 -0.0184 1.0000 0.1180
-5.750 -0.4047 0.08110 0.07696 -0.0236 1.0000 0.1211
-5.500 -0.4045 0.07714 0.07311 -0.0179 1.0000 0.1231
-5.250 -0.4048 0.07492 0.07094 -0.0153 1.0000 0.1256
-4.750 -0.3502 0.06719 0.06301 -0.0266 0.9877 0.1381
-4.500 -0.3044 0.06344 0.05896 -0.0367 0.9773 0.1498
-4.250 -0.2762 0.05926 0.05484 -0.0378 0.9710 0.1554
-4.000 -0.2375 0.05561 0.05094 -0.0445 0.9600 0.1655
-3.750 -0.2051 0.05210 0.04735 -0.0472 0.9514 0.1706
-3.500 -0.1569 0.03735 0.03113 -0.0568 0.9439 0.0942
-3.250 -0.1254 0.03353 0.02684 -0.0581 0.9346 0.0911
-3.000 -0.0879 0.02953 0.02222 -0.0603 0.9283 0.0873
-2.750 -0.0560 0.02642 0.01836 -0.0607 0.9189 0.0848
-2.500 -0.0121 0.02441 0.01586 -0.0633 0.9132 0.0853
-2.250 0.0221 0.02342 0.01463 -0.0643 0.9029 0.0900
-2.000 0.0695 0.02213 0.01308 -0.0675 0.8979 0.0934
-1.750 0.1020 0.02134 0.01234 -0.0682 0.8875 0.0983
-1.500 0.1505 0.02041 0.01143 -0.0717 0.8829 0.1103
-1.250 0.1826 0.02001 0.01124 -0.0722 0.8723 0.1448
-1.000 0.2275 0.02002 0.01120 -0.0749 0.8669 0.2209
-0.750 0.2535 0.02008 0.01125 -0.0744 0.8552 0.2495
-0.500 0.2867 0.01990 0.01104 -0.0752 0.8464 0.2688
-0.250 0.3233 0.01962 0.01078 -0.0766 0.8388 0.2892
0.000 0.3521 0.01956 0.01072 -0.0767 0.8288 0.3072
0.250 0.3926 0.01918 0.01035 -0.0787 0.8227 0.3302
0.500 0.4196 0.01916 0.01038 -0.0785 0.8121 0.3539
0.750 0.4606 0.01868 0.00995 -0.0806 0.8068 0.3877
1.000 0.4841 0.01855 0.00995 -0.0798 0.7957 0.4191
1.250 0.5095 0.01808 0.00974 -0.0793 0.7873 0.4726
1.500 0.6469 0.01654 0.00909 -0.1013 0.7824 1.0000
1.750 0.6728 0.01666 0.00911 -0.1007 0.7733 1.0000
2.000 0.6993 0.01675 0.00910 -0.1000 0.7642 1.0000
2.250 0.7212 0.01695 0.00927 -0.0986 0.7528 1.0000
2.500 0.7457 0.01699 0.00924 -0.0974 0.7409 1.0000
2.750 0.7705 0.01691 0.00907 -0.0961 0.7267 1.0000
3.000 0.7946 0.01681 0.00889 -0.0945 0.7108 1.0000
3.250 0.8182 0.01678 0.00880 -0.0929 0.6950 1.0000
3.500 0.8413 0.01679 0.00876 -0.0913 0.6794 1.0000
3.750 0.8640 0.01685 0.00877 -0.0897 0.6640 1.0000
4.000 0.8865 0.01698 0.00887 -0.0882 0.6499 1.0000
4.250 0.9090 0.01714 0.00905 -0.0868 0.6368 1.0000
4.500 0.9310 0.01726 0.00917 -0.0853 0.6222 1.0000
4.750 0.9525 0.01738 0.00930 -0.0836 0.6070 1.0000
5.000 0.9742 0.01754 0.00948 -0.0821 0.5930 1.0000
5.250 0.9956 0.01770 0.00971 -0.0805 0.5787 1.0000
5.500 1.0163 0.01786 0.00992 -0.0788 0.5637 1.0000
5.750 1.0360 0.01801 0.01014 -0.0769 0.5474 1.0000
6.000 1.0548 0.01814 0.01035 -0.0749 0.5293 1.0000
6.250 1.0735 0.01827 0.01056 -0.0728 0.5104 1.0000
6.500 1.0921 0.01843 0.01078 -0.0707 0.4916 1.0000
6.750 1.1077 0.01863 0.01109 -0.0681 0.4677 1.0000
7.000 1.1239 0.01886 0.01139 -0.0657 0.4439 1.0000
7.250 1.1386 0.01914 0.01172 -0.0629 0.4163 1.0000
7.500 1.1500 0.01953 0.01210 -0.0597 0.3784 1.0000
7.750 1.1550 0.02025 0.01252 -0.0555 0.3236 1.0000
8.000 1.1556 0.02138 0.01327 -0.0508 0.2640 1.0000
8.250 1.1529 0.02284 0.01431 -0.0460 0.2026 1.0000
8.500 1.1475 0.02454 0.01557 -0.0409 0.1411 1.0000
8.750 1.1420 0.02616 0.01685 -0.0358 0.1065 1.0000
9.000 1.1384 0.02777 0.01829 -0.0312 0.0917 1.0000
9.250 1.1361 0.02943 0.01989 -0.0271 0.0820 1.0000
9.500 1.1330 0.03126 0.02176 -0.0233 0.0757 1.0000
9.750 1.1311 0.03324 0.02371 -0.0199 0.0711 1.0000
10.000 1.1371 0.03535 0.02574 -0.0171 0.0675 1.0000
10.250 1.1557 0.03726 0.02765 -0.0156 0.0626 1.0000
10.500 1.2209 0.04136 0.03163 -0.0199 0.0563 1.0000
10.750 1.2511 0.04415 0.03476 -0.0202 0.0546 1.0000
11.000 1.2702 0.04686 0.03778 -0.0192 0.0527 1.0000
11.250 1.2902 0.05004 0.04114 -0.0188 0.0501 1.0000
11.500 1.3003 0.05314 0.04456 -0.0168 0.0498 1.0000
11.750 1.2973 0.05580 0.04759 -0.0131 0.0500 1.0000
12.000 1.2867 0.05842 0.05060 -0.0088 0.0507 1.0000
12.250 1.2731 0.06127 0.05382 -0.0049 0.0516 1.0000
12.500 1.2584 0.06457 0.05745 -0.0018 0.0525 1.0000
12.750 1.2438 0.06820 0.06138 0.0006 0.0535 1.0000
13.000 1.2273 0.07232 0.06577 0.0024 0.0545 1.0000
13.250 1.2124 0.07685 0.07051 0.0034 0.0556 1.0000
13.500 1.0760 0.07598 0.07013 0.0093 0.0547 1.0000
13.750 1.0508 0.08151 0.07587 0.0088 0.0553 1.0000
14.000 1.0238 0.08727 0.08182 0.0076 0.0558 1.0000
14.250 0.9970 0.09328 0.08799 0.0058 0.0565 1.0000
14.500 0.9945 0.09846 0.09328 0.0053 0.0583 1.0000
14.750 0.9524 0.10397 0.09900 0.0019 0.0586 1.0000
15.000 0.8989 0.11152 0.10677 -0.0043 0.0592 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 310 (MVA H.42) AIRFOIL (goe310-il)