GOE 801 (MVA 301) AIRFOIL (goe801-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 801 (MVA 301) AIRFOIL (goe801-il) Reynolds number: 200,000 Max Cl/Cd: 79.55 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe801-il-200000-n5.txt Download as CSV file: xf-goe801-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 801 (MVA 301) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1489 0.09242 0.08884 -0.0591 0.9596 0.0299
-8.000 -0.1369 0.08945 0.08586 -0.0616 0.9493 0.0310
-7.750 -0.1259 0.08643 0.08283 -0.0644 0.9396 0.0320
-7.500 -0.1184 0.08354 0.07993 -0.0688 0.9284 0.0328
-7.250 -0.1068 0.08017 0.07653 -0.0763 0.9162 0.0331
-7.000 -0.0891 0.07629 0.07259 -0.0834 0.9072 0.0332
-6.750 -0.0801 0.07335 0.06967 -0.0810 0.9004 0.0335
-6.500 -0.0676 0.07085 0.06715 -0.0801 0.8943 0.0341
-6.250 -0.0530 0.06835 0.06464 -0.0814 0.8858 0.0349
-6.000 -0.0352 0.06550 0.06173 -0.0843 0.8791 0.0362
-5.750 -0.0143 0.06230 0.05848 -0.0888 0.8703 0.0376
-5.500 0.0294 0.05694 0.05282 -0.1028 0.8629 0.0393
-5.250 0.0470 0.05353 0.04939 -0.1042 0.8549 0.0396
-5.000 0.0630 0.05117 0.04702 -0.1039 0.8483 0.0401
-4.750 0.0830 0.04893 0.04473 -0.1048 0.8406 0.0409
-4.500 0.1066 0.04662 0.04234 -0.1066 0.8332 0.0426
-4.250 0.1530 0.04194 0.03717 -0.1147 0.8259 0.0468
-4.000 0.1717 0.03958 0.03483 -0.1148 0.8179 0.0474
-3.750 0.1942 0.03768 0.03287 -0.1153 0.8100 0.0483
-3.500 0.2199 0.03572 0.03081 -0.1163 0.8012 0.0498
-3.250 0.2590 0.03277 0.02725 -0.1195 0.7926 0.0555
-3.000 0.2817 0.03059 0.02508 -0.1198 0.7828 0.0565
-2.750 0.3064 0.02914 0.02357 -0.1201 0.7713 0.0581
-2.500 0.3379 0.02751 0.02151 -0.1211 0.7607 0.0664
-2.250 0.3730 0.02273 0.01597 -0.1214 0.7512 0.0390
-2.000 0.4002 0.02128 0.01438 -0.1216 0.7400 0.0382
-1.750 0.4284 0.01996 0.01280 -0.1217 0.7291 0.0375
-1.500 0.4569 0.01881 0.01135 -0.1216 0.7176 0.0370
-1.250 0.4852 0.01784 0.01012 -0.1215 0.7051 0.0368
-1.000 0.5134 0.01715 0.00918 -0.1213 0.6926 0.0377
-0.750 0.5414 0.01657 0.00835 -0.1211 0.6800 0.0383
-0.500 0.5691 0.01598 0.00756 -0.1209 0.6670 0.0382
-0.250 0.5965 0.01548 0.00690 -0.1206 0.6533 0.0381
0.000 0.6239 0.01506 0.00635 -0.1203 0.6396 0.0382
0.250 0.6512 0.01472 0.00590 -0.1200 0.6266 0.0384
0.500 0.6782 0.01445 0.00552 -0.1197 0.6138 0.0386
0.750 0.7051 0.01424 0.00523 -0.1193 0.6010 0.0390
1.000 0.7318 0.01405 0.00498 -0.1190 0.5882 0.0396
1.250 0.7584 0.01385 0.00479 -0.1188 0.5757 0.0414
1.500 0.7852 0.01378 0.00469 -0.1186 0.5643 0.0431
1.750 0.8119 0.01375 0.00460 -0.1183 0.5535 0.0441
2.000 0.8387 0.01376 0.00456 -0.1181 0.5422 0.0451
2.250 0.8653 0.01381 0.00454 -0.1178 0.5311 0.0463
2.500 0.8917 0.01388 0.00454 -0.1175 0.5206 0.0484
2.750 0.9182 0.01396 0.00459 -0.1172 0.5102 0.0524
3.250 0.9669 0.01253 0.00494 -0.1160 0.4915 1.0000
3.500 0.9930 0.01275 0.00507 -0.1157 0.4817 1.0000
3.750 1.0185 0.01301 0.00521 -0.1152 0.4716 1.0000
4.000 1.0439 0.01326 0.00537 -0.1148 0.4598 1.0000
4.250 1.0691 0.01352 0.00556 -0.1143 0.4478 1.0000
4.500 1.0938 0.01380 0.00576 -0.1138 0.4357 1.0000
4.750 1.1182 0.01410 0.00597 -0.1132 0.4238 1.0000
5.000 1.1426 0.01438 0.00621 -0.1127 0.4116 1.0000
5.250 1.1670 0.01467 0.00646 -0.1122 0.4006 1.0000
5.500 1.1909 0.01499 0.00672 -0.1116 0.3910 1.0000
5.750 1.2153 0.01529 0.00702 -0.1111 0.3823 1.0000
6.000 1.2391 0.01562 0.00732 -0.1105 0.3747 1.0000
6.250 1.2628 0.01594 0.00764 -0.1099 0.3657 1.0000
6.500 1.2857 0.01630 0.00799 -0.1091 0.3561 1.0000
6.750 1.3077 0.01669 0.00835 -0.1083 0.3456 1.0000
7.000 1.3301 0.01705 0.00872 -0.1075 0.3349 1.0000
7.250 1.3517 0.01746 0.00911 -0.1066 0.3255 1.0000
7.500 1.3731 0.01786 0.00954 -0.1057 0.3159 1.0000
7.750 1.3943 0.01827 0.00997 -0.1047 0.3065 1.0000
8.000 1.4137 0.01875 0.01044 -0.1035 0.2965 1.0000
8.250 1.4334 0.01919 0.01093 -0.1024 0.2844 1.0000
8.500 1.4518 0.01969 0.01143 -0.1010 0.2715 1.0000
8.750 1.4691 0.02023 0.01198 -0.0995 0.2588 1.0000
9.000 1.4837 0.02081 0.01257 -0.0976 0.2446 1.0000
9.250 1.4952 0.02154 0.01326 -0.0953 0.2261 1.0000
9.500 1.5042 0.02247 0.01409 -0.0928 0.2063 1.0000
9.750 1.5128 0.02350 0.01504 -0.0903 0.1885 1.0000
10.000 1.5211 0.02460 0.01609 -0.0880 0.1747 1.0000
10.500 1.5396 0.02683 0.01834 -0.0840 0.1552 1.0000
10.750 1.5479 0.02806 0.01959 -0.0821 0.1475 1.0000
11.000 1.5571 0.02927 0.02085 -0.0803 0.1393 1.0000
11.250 1.5647 0.03064 0.02224 -0.0786 0.1313 1.0000
11.500 1.5715 0.03212 0.02376 -0.0769 0.1229 1.0000
11.750 1.5792 0.03356 0.02526 -0.0754 0.1159 1.0000
12.000 1.5846 0.03524 0.02698 -0.0738 0.1086 1.0000
12.250 1.5908 0.03690 0.02870 -0.0725 0.1010 1.0000
12.500 1.5943 0.03886 0.03069 -0.0711 0.0934 1.0000
12.750 1.5974 0.04092 0.03279 -0.0698 0.0849 1.0000
13.000 1.5990 0.04318 0.03509 -0.0686 0.0774 1.0000
13.250 1.5991 0.04567 0.03762 -0.0675 0.0715 1.0000
13.500 1.5975 0.04841 0.04039 -0.0666 0.0655 1.0000
13.750 1.5948 0.05136 0.04340 -0.0658 0.0604 1.0000
14.000 1.5925 0.05440 0.04651 -0.0652 0.0557 1.0000
14.250 1.5890 0.05768 0.04986 -0.0647 0.0514 1.0000
14.500 1.5866 0.06094 0.05322 -0.0645 0.0456 1.0000
14.750 1.5825 0.06451 0.05687 -0.0644 0.0368 1.0000
15.000 1.5639 0.07008 0.06236 -0.0647 0.0224 1.0000
15.250 1.5512 0.07506 0.06739 -0.0652 0.0206 1.0000
15.500 1.5404 0.07993 0.07237 -0.0658 0.0192 1.0000
15.750 1.5318 0.08458 0.07716 -0.0666 0.0185 1.0000
16.000 1.5230 0.08933 0.08205 -0.0675 0.0179 1.0000
16.250 1.5137 0.09426 0.08712 -0.0685 0.0174 1.0000
16.500 1.5038 0.09936 0.09237 -0.0698 0.0170 1.0000
16.750 1.4933 0.10466 0.09783 -0.0712 0.0167 1.0000
17.000 1.4821 0.11015 0.10347 -0.0729 0.0163 1.0000
17.250 1.4705 0.11581 0.10928 -0.0748 0.0161 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 801 (MVA 301) AIRFOIL (goe801-il)