GOE 225 (MVA H.35) AIRFOIL (goe225-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 225 (MVA H.35) AIRFOIL (goe225-il) Reynolds number: 200,000 Max Cl/Cd: 83.26 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe225-il-200000-n5.txt Download as CSV file: xf-goe225-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 225 (MVA H.35) 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.0004 0.08429 0.07999 -0.1069 0.8828 0.0382
-7.750 0.0037 0.02764 0.02199 -0.1885 0.8614 0.0376
-7.500 0.0419 0.02491 0.01881 -0.1944 0.8588 0.0382
-7.250 0.0757 0.02319 0.01677 -0.1977 0.8557 0.0388
-7.000 0.1077 0.02193 0.01525 -0.2000 0.8520 0.0395
-6.750 0.1396 0.02090 0.01398 -0.2017 0.8481 0.0402
-6.500 0.1705 0.02002 0.01303 -0.2030 0.8444 0.0411
-6.250 0.2021 0.01928 0.01219 -0.2041 0.8412 0.0418
-6.000 0.2327 0.01865 0.01146 -0.2051 0.8375 0.0426
-5.750 0.2629 0.01809 0.01082 -0.2060 0.8337 0.0434
-5.500 0.2937 0.01756 0.01020 -0.2069 0.8300 0.0444
-5.250 0.3249 0.01707 0.00959 -0.2077 0.8268 0.0455
-5.000 0.3564 0.01659 0.00904 -0.2086 0.8240 0.0467
-4.750 0.3873 0.01622 0.00863 -0.2094 0.8208 0.0485
-4.500 0.4168 0.01593 0.00831 -0.2099 0.8162 0.0511
-4.250 0.4471 0.01561 0.00797 -0.2104 0.8118 0.0544
-4.000 0.4779 0.01531 0.00763 -0.2110 0.8079 0.0591
-3.750 0.5091 0.01504 0.00733 -0.2116 0.8047 0.0669
-3.500 0.5375 0.01493 0.00723 -0.2117 0.7992 0.0776
-3.250 0.5666 0.01482 0.00707 -0.2118 0.7938 0.0878
-3.000 0.5966 0.01464 0.00686 -0.2121 0.7891 0.0969
-2.750 0.6258 0.01451 0.00670 -0.2124 0.7839 0.1044
-2.500 0.6542 0.01440 0.00657 -0.2124 0.7772 0.1122
-2.250 0.6841 0.01420 0.00633 -0.2127 0.7715 0.1194
-2.000 0.7126 0.01407 0.00621 -0.2128 0.7640 0.1272
-1.750 0.7416 0.01389 0.00605 -0.2130 0.7563 0.1373
-1.500 0.7704 0.01375 0.00591 -0.2131 0.7482 0.1485
-1.250 0.7993 0.01357 0.00577 -0.2132 0.7386 0.1623
-1.000 0.8274 0.01343 0.00572 -0.2133 0.7272 0.1875
-0.750 0.8556 0.01331 0.00568 -0.2132 0.7146 0.2297
-0.500 0.8835 0.01327 0.00562 -0.2130 0.6999 0.2613
-0.250 0.9111 0.01324 0.00552 -0.2127 0.6823 0.2821
0.000 0.9384 0.01325 0.00541 -0.2124 0.6628 0.2979
0.250 0.9652 0.01331 0.00535 -0.2119 0.6435 0.3113
0.500 0.9914 0.01344 0.00537 -0.2114 0.6261 0.3254
0.750 1.0173 0.01362 0.00545 -0.2109 0.6105 0.3405
1.000 1.0432 0.01383 0.00555 -0.2105 0.5969 0.3534
1.250 1.0690 0.01404 0.00571 -0.2100 0.5855 0.3666
1.500 1.0949 0.01426 0.00589 -0.2096 0.5753 0.3795
1.750 1.1210 0.01448 0.00608 -0.2092 0.5663 0.3942
2.000 1.1468 0.01472 0.00630 -0.2088 0.5578 0.4113
2.250 1.1727 0.01496 0.00653 -0.2085 0.5504 0.4292
2.500 1.1985 0.01519 0.00677 -0.2081 0.5430 0.4473
2.750 1.2241 0.01547 0.00702 -0.2076 0.5364 0.4662
3.000 1.2498 0.01568 0.00727 -0.2072 0.5296 0.4846
3.250 1.2748 0.01595 0.00753 -0.2067 0.5220 0.5009
3.500 1.2993 0.01621 0.00780 -0.2061 0.5136 0.5166
3.750 1.3232 0.01649 0.00806 -0.2053 0.5046 0.5322
4.000 1.3474 0.01676 0.00834 -0.2047 0.4972 0.5483
4.250 1.3713 0.01699 0.00862 -0.2040 0.4901 0.5653
4.750 1.4185 0.01749 0.00920 -0.2024 0.4774 0.6011
5.000 1.4416 0.01772 0.00950 -0.2016 0.4708 0.6223
5.250 1.4643 0.01798 0.00981 -0.2007 0.4650 0.6532
5.500 1.4812 0.01779 0.01005 -0.1985 0.4589 0.9232
5.750 1.5033 0.01811 0.01037 -0.1975 0.4525 1.0000
6.000 1.5248 0.01848 0.01072 -0.1964 0.4462 1.0000
6.250 1.5457 0.01881 0.01110 -0.1952 0.4389 1.0000
6.500 1.5652 0.01921 0.01148 -0.1938 0.4322 1.0000
6.750 1.5850 0.01958 0.01191 -0.1924 0.4254 1.0000
7.000 1.6024 0.01998 0.01233 -0.1906 0.4182 1.0000
7.250 1.6184 0.02043 0.01280 -0.1885 0.4110 1.0000
7.500 1.6335 0.02092 0.01332 -0.1864 0.4017 1.0000
7.750 1.6479 0.02147 0.01390 -0.1842 0.3918 1.0000
8.000 1.6598 0.02215 0.01454 -0.1817 0.3805 1.0000
8.250 1.6725 0.02283 0.01524 -0.1795 0.3678 1.0000
8.500 1.6848 0.02358 0.01601 -0.1772 0.3554 1.0000
8.750 1.6959 0.02444 0.01686 -0.1748 0.3432 1.0000
9.000 1.7062 0.02538 0.01779 -0.1725 0.3312 1.0000
9.250 1.7165 0.02637 0.01877 -0.1702 0.3184 1.0000
9.500 1.7263 0.02743 0.01984 -0.1679 0.3057 1.0000
9.750 1.7348 0.02861 0.02102 -0.1656 0.2933 1.0000
10.000 1.7421 0.02991 0.02230 -0.1632 0.2814 1.0000
10.250 1.7486 0.03131 0.02369 -0.1608 0.2696 1.0000
10.500 1.7563 0.03267 0.02507 -0.1586 0.2587 1.0000
10.750 1.7620 0.03421 0.02660 -0.1563 0.2484 1.0000
11.000 1.7669 0.03584 0.02824 -0.1541 0.2386 1.0000
11.250 1.7735 0.03738 0.02982 -0.1520 0.2300 1.0000
11.500 1.7761 0.03927 0.03170 -0.1497 0.2212 1.0000
11.750 1.7815 0.04096 0.03345 -0.1477 0.2119 1.0000
12.000 1.7834 0.04300 0.03551 -0.1456 0.2034 1.0000
12.250 1.7867 0.04500 0.03756 -0.1437 0.1942 1.0000
12.500 1.7891 0.04714 0.03974 -0.1418 0.1860 1.0000
12.750 1.7902 0.04946 0.04209 -0.1400 0.1776 1.0000
13.000 1.7926 0.05172 0.04442 -0.1384 0.1693 1.0000
13.250 1.7906 0.05452 0.04725 -0.1367 0.1599 1.0000
13.500 1.7889 0.05741 0.05016 -0.1352 0.1485 1.0000
13.750 1.7860 0.06050 0.05330 -0.1338 0.1364 1.0000
14.000 1.7792 0.06415 0.05695 -0.1326 0.1218 1.0000
14.250 1.7675 0.06853 0.06129 -0.1314 0.1021 1.0000
14.500 1.7438 0.07462 0.06721 -0.1304 0.0750 1.0000
14.750 1.7274 0.07992 0.07248 -0.1298 0.0654 1.0000
15.000 1.7160 0.08469 0.07729 -0.1294 0.0604 1.0000
15.250 1.7062 0.08935 0.08204 -0.1292 0.0567 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 225 (MVA H.35) AIRFOIL (goe225-il)