GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Reynolds number: 500,000 Max Cl/Cd: 96.89 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe207-il-500000-n5.txt Download as CSV file: xf-goe207-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 207 (AVIATIK V8) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3042 0.11823 0.11578 -0.0017 0.8535 0.0173
-9.500 -0.2984 0.11540 0.11280 -0.0028 0.8168 0.0173
-9.250 -0.2920 0.11253 0.10979 -0.0042 0.7850 0.0173
-9.000 -0.2853 0.10963 0.10677 -0.0058 0.7580 0.0174
-8.750 -0.2778 0.10662 0.10366 -0.0070 0.7349 0.0174
-8.500 -0.2690 0.10382 0.10078 -0.0075 0.7136 0.0175
-8.250 -0.2600 0.10129 0.09817 -0.0085 0.6941 0.0178
-8.000 -0.2508 0.09878 0.09559 -0.0097 0.6753 0.0181
-7.750 -0.2416 0.09619 0.09294 -0.0112 0.6571 0.0184
-7.500 -0.2326 0.09354 0.09021 -0.0128 0.6391 0.0188
-7.250 -0.2243 0.09093 0.08754 -0.0145 0.6218 0.0193
-7.000 -0.2119 0.08797 0.08452 -0.0176 0.6054 0.0198
-6.750 -0.1965 0.08481 0.08128 -0.0217 0.5906 0.0201
-6.500 -0.1787 0.08158 0.07798 -0.0264 0.5773 0.0201
-6.250 -0.1569 0.07819 0.07450 -0.0325 0.5653 0.0202
-6.000 -0.1339 0.07477 0.07099 -0.0381 0.5540 0.0203
-5.750 -0.1110 0.07134 0.06747 -0.0427 0.5441 0.0203
-5.250 -0.0767 0.06558 0.06161 -0.0453 0.5250 0.0206
-5.000 -0.0545 0.06294 0.05889 -0.0480 0.5157 0.0208
-4.750 -0.0300 0.06029 0.05618 -0.0512 0.5065 0.0211
-4.500 -0.0039 0.05766 0.05345 -0.0545 0.4978 0.0215
-4.250 0.0239 0.05500 0.05070 -0.0579 0.4892 0.0221
-4.000 0.0596 0.05215 0.04771 -0.0627 0.4817 0.0234
-3.750 0.0958 0.04925 0.04466 -0.0673 0.4746 0.0235
-3.250 0.1639 0.04382 0.03891 -0.0739 0.4595 0.0236
-3.000 0.1955 0.04136 0.03630 -0.0762 0.4509 0.0237
-2.750 0.2187 0.03891 0.03380 -0.0775 0.4430 0.0239
-2.500 0.2447 0.03712 0.03193 -0.0787 0.4343 0.0241
-2.250 0.2733 0.03548 0.03021 -0.0802 0.4264 0.0246
-2.000 0.3041 0.03380 0.02839 -0.0818 0.4176 0.0253
-1.750 0.3361 0.03206 0.02651 -0.0833 0.4094 0.0259
-1.500 0.3747 0.03033 0.02451 -0.0849 0.4005 0.0273
-1.250 0.4088 0.02859 0.02256 -0.0861 0.3923 0.0275
-1.000 0.4411 0.02695 0.02071 -0.0871 0.3835 0.0275
-0.750 0.4726 0.02544 0.01903 -0.0879 0.3753 0.0276
-0.500 0.5005 0.02388 0.01736 -0.0892 0.3664 0.0280
-0.250 0.5288 0.02307 0.01648 -0.0899 0.3578 0.0288
0.000 0.5589 0.02220 0.01546 -0.0905 0.3493 0.0300
0.250 0.5904 0.02113 0.01422 -0.0911 0.3417 0.0308
0.750 0.6552 0.01825 0.01086 -0.0920 0.3288 0.0260
1.000 0.6863 0.01714 0.00953 -0.0923 0.3226 0.0259
1.250 0.7169 0.01611 0.00828 -0.0925 0.3173 0.0260
1.500 0.7473 0.01510 0.00706 -0.0927 0.3123 0.0265
1.750 0.7768 0.01433 0.00608 -0.0927 0.3070 0.0272
2.000 0.8056 0.01389 0.00548 -0.0928 0.3015 0.0272
2.250 0.8343 0.01357 0.00507 -0.0929 0.2961 0.0275
2.500 0.8627 0.01330 0.00472 -0.0931 0.2903 0.0279
2.750 0.8910 0.01318 0.00456 -0.0932 0.2851 0.0284
3.000 0.9194 0.01310 0.00448 -0.0933 0.2809 0.0291
3.250 0.9477 0.01307 0.00444 -0.0935 0.2769 0.0298
3.500 0.9758 0.01312 0.00445 -0.0936 0.2725 0.0311
3.750 1.0039 0.01317 0.00448 -0.0937 0.2686 0.0321
4.000 1.0322 0.01315 0.00449 -0.0938 0.2648 0.0330
4.250 1.0603 0.01322 0.00456 -0.0939 0.2613 0.0343
4.500 1.0881 0.01333 0.00466 -0.0940 0.2581 0.0359
4.750 1.1157 0.01347 0.00479 -0.0941 0.2550 0.0380
5.000 1.1434 0.01358 0.00494 -0.0941 0.2522 0.0451
5.250 1.1711 0.01369 0.00512 -0.0942 0.2477 0.0713
5.500 1.1983 0.01386 0.00530 -0.0942 0.2427 0.0950
6.000 1.2465 0.01293 0.00579 -0.0932 0.2345 1.0000
6.250 1.2734 0.01315 0.00598 -0.0932 0.2295 1.0000
6.500 1.2998 0.01344 0.00622 -0.0931 0.2239 1.0000
6.750 1.3264 0.01369 0.00645 -0.0931 0.2184 1.0000
7.000 1.3525 0.01399 0.00670 -0.0930 0.2102 1.0000
7.250 1.3783 0.01431 0.00700 -0.0929 0.2021 1.0000
7.500 1.4034 0.01470 0.00732 -0.0927 0.1904 1.0000
7.750 1.4187 0.01639 0.00850 -0.0916 0.1158 1.0000
8.000 1.4388 0.01736 0.00938 -0.0909 0.0948 1.0000
8.250 1.4502 0.01931 0.01102 -0.0892 0.0318 1.0000
8.500 1.4717 0.01994 0.01168 -0.0885 0.0272 1.0000
8.750 1.4928 0.02058 0.01236 -0.0878 0.0249 1.0000
9.000 1.5134 0.02121 0.01306 -0.0871 0.0231 1.0000
9.250 1.5337 0.02181 0.01372 -0.0863 0.0216 1.0000
9.500 1.5526 0.02251 0.01447 -0.0853 0.0204 1.0000
9.750 1.5690 0.02337 0.01539 -0.0841 0.0193 1.0000
10.000 1.5860 0.02409 0.01619 -0.0830 0.0186 1.0000
10.250 1.6007 0.02490 0.01708 -0.0815 0.0178 1.0000
10.500 1.6127 0.02580 0.01804 -0.0797 0.0170 1.0000
10.750 1.6195 0.02688 0.01918 -0.0773 0.0164 1.0000
11.000 1.6240 0.02829 0.02069 -0.0751 0.0158 1.0000
11.250 1.6261 0.03014 0.02264 -0.0733 0.0154 1.0000
11.500 1.6307 0.03202 0.02462 -0.0723 0.0151 1.0000
11.750 1.6338 0.03427 0.02698 -0.0716 0.0147 1.0000
12.000 1.6349 0.03695 0.02978 -0.0714 0.0144 1.0000
12.250 1.6339 0.04009 0.03305 -0.0715 0.0141 1.0000
12.500 1.6306 0.04370 0.03677 -0.0719 0.0138 1.0000
12.750 1.6248 0.04778 0.04100 -0.0725 0.0136 1.0000
13.000 1.6161 0.05237 0.04571 -0.0734 0.0134 1.0000
13.250 1.6046 0.05741 0.05089 -0.0744 0.0132 1.0000
13.500 1.5901 0.06287 0.05649 -0.0755 0.0131 1.0000
13.750 1.5734 0.06864 0.06239 -0.0766 0.0130 1.0000
14.000 1.5549 0.07470 0.06858 -0.0777 0.0129 1.0000
14.250 1.5361 0.08093 0.07495 -0.0791 0.0128 1.0000
14.500 1.5180 0.08725 0.08140 -0.0805 0.0127 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)