GOE 409 AIRFOIL (goe409-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 409 AIRFOIL (goe409-il) Reynolds number: 100,000 Max Cl/Cd: 38.57 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe409-il-100000.txt Download as CSV file: xf-goe409-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 409 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7858 0.07787 0.07225 -0.0482 1.0000 0.1022
-11.750 -0.8050 0.07247 0.06678 -0.0488 1.0000 0.1015
-11.500 -0.8306 0.06773 0.06194 -0.0481 1.0000 0.1008
-11.250 -0.8544 0.06362 0.05770 -0.0460 1.0000 0.0997
-11.000 -0.8823 0.06020 0.05412 -0.0420 1.0000 0.0991
-10.750 -0.9039 0.05672 0.05042 -0.0379 1.0000 0.0985
-10.500 -0.9213 0.05339 0.04683 -0.0337 1.0000 0.0986
-10.250 -0.9338 0.05028 0.04341 -0.0295 1.0000 0.0991
-10.000 -0.9429 0.04746 0.04025 -0.0252 1.0000 0.0999
-9.750 -0.9488 0.04493 0.03733 -0.0209 1.0000 0.1010
-9.500 -0.9513 0.04268 0.03467 -0.0167 1.0000 0.1021
-9.250 -0.9485 0.04021 0.03187 -0.0133 1.0000 0.1035
-9.000 -0.9351 0.03802 0.02962 -0.0115 1.0000 0.1055
-8.750 -0.9216 0.03631 0.02777 -0.0093 1.0000 0.1073
-8.500 -0.9083 0.03479 0.02608 -0.0071 1.0000 0.1094
-8.250 -0.8949 0.03320 0.02425 -0.0047 1.0000 0.1111
-8.000 -0.8793 0.03160 0.02238 -0.0027 1.0000 0.1122
-7.750 -0.8622 0.03016 0.02071 -0.0008 1.0000 0.1135
-7.500 -0.8450 0.02901 0.01930 0.0012 1.0000 0.1154
-7.250 -0.8232 0.02755 0.01780 0.0021 1.0000 0.1174
-7.000 -0.8016 0.02639 0.01666 0.0032 1.0000 0.1195
-6.750 -0.7807 0.02540 0.01564 0.0045 1.0000 0.1216
-6.500 -0.7605 0.02450 0.01471 0.0059 1.0000 0.1242
-6.250 -0.7409 0.02371 0.01385 0.0074 1.0000 0.1271
-6.000 -0.7224 0.02292 0.01306 0.0091 1.0000 0.1309
-5.750 -0.7058 0.02219 0.01242 0.0110 1.0000 0.1354
-5.500 -0.6892 0.02156 0.01179 0.0130 1.0000 0.1407
-5.250 -0.6742 0.02087 0.01116 0.0153 1.0000 0.1468
-5.000 -0.6599 0.02024 0.01059 0.0176 1.0000 0.1557
-4.750 -0.6469 0.01951 0.00998 0.0202 1.0000 0.1703
-4.500 -0.6361 0.01856 0.00932 0.0230 1.0000 0.2051
-4.250 -0.6304 0.01715 0.00869 0.0264 1.0000 0.3054
-4.000 -0.6246 0.01621 0.00856 0.0302 1.0000 0.4506
-3.750 -0.6118 0.01587 0.00847 0.0331 1.0000 0.5257
-3.500 -0.5974 0.01563 0.00848 0.0359 1.0000 0.5815
-3.250 -0.5810 0.01551 0.00851 0.0382 1.0000 0.6243
-3.000 -0.5644 0.01547 0.00862 0.0406 1.0000 0.6668
-2.750 -0.5470 0.01550 0.00874 0.0428 1.0000 0.7038
-2.500 -0.5286 0.01563 0.00898 0.0450 1.0000 0.7402
-2.250 -0.5099 0.01588 0.00933 0.0472 1.0000 0.7769
-2.000 -0.4898 0.01624 0.00974 0.0491 1.0000 0.8109
-1.750 -0.4670 0.01669 0.01021 0.0504 1.0000 0.8394
-1.500 -0.4372 0.01728 0.01077 0.0502 0.9982 0.8660
-1.250 -0.3816 0.01848 0.01190 0.0455 0.9902 0.8940
-1.000 -0.3056 0.02009 0.01344 0.0379 0.9861 0.9171
-0.750 -0.2156 0.02144 0.01470 0.0271 0.9842 0.9339
-0.500 -0.1412 0.02200 0.01520 0.0179 0.9785 0.9458
-0.250 -0.0714 0.02231 0.01547 0.0092 0.9717 0.9568
0.000 0.0001 0.02233 0.01548 0.0000 0.9636 0.9636
0.250 0.0713 0.02231 0.01547 -0.0092 0.9568 0.9717
0.500 0.1414 0.02199 0.01519 -0.0179 0.9458 0.9784
0.750 0.2153 0.02144 0.01471 -0.0271 0.9340 0.9842
1.000 0.3053 0.02010 0.01345 -0.0379 0.9173 0.9862
1.250 0.3815 0.01848 0.01189 -0.0455 0.8939 0.9903
1.500 0.4372 0.01727 0.01076 -0.0502 0.8661 0.9983
1.750 0.4669 0.01669 0.01020 -0.0504 0.8396 1.0000
2.000 0.4897 0.01624 0.00974 -0.0491 0.8108 1.0000
2.250 0.5097 0.01588 0.00932 -0.0472 0.7768 1.0000
2.500 0.5285 0.01563 0.00899 -0.0450 0.7406 1.0000
2.750 0.5468 0.01550 0.00874 -0.0428 0.7040 1.0000
3.000 0.5643 0.01547 0.00862 -0.0406 0.6674 1.0000
3.250 0.5808 0.01551 0.00850 -0.0382 0.6242 1.0000
3.500 0.5972 0.01562 0.00848 -0.0358 0.5821 1.0000
3.750 0.6117 0.01586 0.00846 -0.0331 0.5263 1.0000
4.000 0.6245 0.01620 0.00856 -0.0302 0.4514 1.0000
4.250 0.6304 0.01713 0.00868 -0.0264 0.3074 1.0000
4.500 0.6360 0.01855 0.00931 -0.0229 0.2054 1.0000
4.750 0.6468 0.01951 0.00997 -0.0201 0.1704 1.0000
5.000 0.6597 0.02024 0.01058 -0.0176 0.1556 1.0000
5.250 0.6740 0.02087 0.01115 -0.0152 0.1467 1.0000
5.500 0.6891 0.02156 0.01179 -0.0130 0.1408 1.0000
5.750 0.7056 0.02218 0.01241 -0.0110 0.1355 1.0000
6.000 0.7222 0.02291 0.01304 -0.0091 0.1309 1.0000
6.250 0.7408 0.02371 0.01385 -0.0074 0.1272 1.0000
6.500 0.7603 0.02450 0.01471 -0.0058 0.1241 1.0000
6.750 0.7806 0.02539 0.01564 -0.0044 0.1216 1.0000
7.000 0.8014 0.02638 0.01665 -0.0032 0.1195 1.0000
7.250 0.8230 0.02753 0.01779 -0.0021 0.1175 1.0000
7.500 0.8449 0.02901 0.01929 -0.0012 0.1154 1.0000
7.750 0.8620 0.03015 0.02069 0.0008 0.1136 1.0000
8.000 0.8792 0.03159 0.02237 0.0027 0.1122 1.0000
8.250 0.8948 0.03319 0.02424 0.0048 0.1111 1.0000
8.500 0.9082 0.03476 0.02605 0.0071 0.1094 1.0000
8.750 0.9215 0.03632 0.02778 0.0094 0.1073 1.0000
9.000 0.9350 0.03802 0.02962 0.0115 0.1056 1.0000
9.250 0.9484 0.04021 0.03187 0.0133 0.1035 1.0000
9.500 0.9509 0.04265 0.03465 0.0168 0.1020 1.0000
9.750 0.9487 0.04492 0.03733 0.0209 0.1010 1.0000
10.000 0.9430 0.04748 0.04027 0.0252 0.1000 1.0000
10.250 0.9338 0.05030 0.04343 0.0295 0.0991 1.0000
10.500 0.9212 0.05336 0.04680 0.0338 0.0985 1.0000
10.750 0.9042 0.05675 0.05046 0.0379 0.0986 1.0000
11.000 0.8824 0.06022 0.05415 0.0419 0.0991 1.0000
11.250 0.8548 0.06373 0.05781 0.0459 0.0998 1.0000
11.500 0.8275 0.06782 0.06204 0.0481 0.1006 1.0000
11.750 0.8021 0.07261 0.06694 0.0487 0.1013 1.0000
12.000 0.7880 0.07791 0.07229 0.0481 0.1022 1.0000
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Polar data table (+)
Polar graphs
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