GOE 795 AIRFOIL (goe795-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 795 AIRFOIL (goe795-il) Reynolds number: 500,000 Max Cl/Cd: 80.16 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe795-il-500000-n5.txt Download as CSV file: xf-goe795-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 795 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5034 0.09160 0.08930 -0.0263 1.0000 0.0085
-9.000 -0.7215 0.03369 0.03090 -0.0477 1.0000 0.0081
-8.750 -0.6994 0.03121 0.02819 -0.0491 0.9982 0.0085
-8.500 -0.6733 0.02941 0.02618 -0.0507 0.9959 0.0091
-8.250 -0.6491 0.02689 0.02333 -0.0521 0.9932 0.0096
-8.000 -0.6291 0.02383 0.01983 -0.0524 0.9895 0.0098
-7.750 -0.6034 0.02212 0.01783 -0.0531 0.9863 0.0102
-7.500 -0.5727 0.02178 0.01740 -0.0542 0.9841 0.0110
-7.250 -0.5444 0.02018 0.01552 -0.0551 0.9819 0.0112
-7.000 -0.5184 0.01994 0.01519 -0.0550 0.9775 0.0117
-6.750 -0.4965 0.01720 0.01196 -0.0548 0.9731 0.0128
-6.500 -0.4676 0.01596 0.01051 -0.0554 0.9704 0.0129
-6.250 -0.4410 0.01493 0.00929 -0.0554 0.9664 0.0129
-6.000 -0.4146 0.01406 0.00827 -0.0552 0.9617 0.0129
-5.750 -0.3853 0.01326 0.00732 -0.0556 0.9584 0.0130
-5.500 -0.3541 0.01250 0.00644 -0.0565 0.9560 0.0132
-5.250 -0.3275 0.01191 0.00575 -0.0562 0.9502 0.0134
-5.000 -0.2961 0.01135 0.00507 -0.0570 0.9464 0.0137
-4.750 -0.2618 0.01086 0.00450 -0.0585 0.9436 0.0141
-4.500 -0.2308 0.01044 0.00402 -0.0592 0.9385 0.0148
-4.250 -0.1985 0.01008 0.00359 -0.0602 0.9331 0.0156
-4.000 -0.1635 0.00975 0.00321 -0.0617 0.9287 0.0167
-3.750 -0.1336 0.00947 0.00288 -0.0621 0.9205 0.0185
-3.500 -0.1009 0.00921 0.00262 -0.0632 0.9135 0.0243
-3.250 -0.0711 0.00900 0.00238 -0.0635 0.9039 0.0300
-3.000 -0.0417 0.00882 0.00217 -0.0638 0.8939 0.0355
-2.750 -0.0124 0.00867 0.00197 -0.0641 0.8831 0.0424
-2.500 0.0156 0.00848 0.00180 -0.0641 0.8700 0.0540
-2.250 0.0426 0.00830 0.00168 -0.0639 0.8543 0.0803
-2.000 0.0694 0.00821 0.00157 -0.0636 0.8378 0.0947
-1.750 0.0956 0.00812 0.00145 -0.0631 0.8202 0.1117
-1.500 0.1210 0.00800 0.00135 -0.0625 0.8025 0.1344
-1.250 0.1459 0.00790 0.00128 -0.0619 0.7840 0.1649
-1.000 0.1708 0.00784 0.00122 -0.0612 0.7660 0.1979
-0.500 0.2201 0.00764 0.00113 -0.0598 0.7369 0.2746
-0.250 0.2424 0.00726 0.00112 -0.0588 0.7262 0.4031
0.000 0.2623 0.00678 0.00113 -0.0572 0.7159 0.5678
0.250 0.2815 0.00642 0.00115 -0.0553 0.7060 0.6970
0.500 0.3067 0.00595 0.00122 -0.0543 0.6947 0.8726
0.750 0.4015 0.00607 0.00132 -0.0694 0.6646 0.9531
1.000 0.4351 0.00623 0.00139 -0.0705 0.6446 0.9764
1.250 0.4705 0.00641 0.00146 -0.0722 0.6209 0.9921
1.500 0.5147 0.00656 0.00150 -0.0759 0.5921 1.0000
1.750 0.5365 0.00672 0.00154 -0.0746 0.5621 1.0000
2.000 0.5571 0.00695 0.00159 -0.0731 0.5192 1.0000
2.250 0.5764 0.00730 0.00168 -0.0713 0.4619 1.0000
2.500 0.5940 0.00782 0.00184 -0.0693 0.3869 1.0000
2.750 0.6137 0.00824 0.00202 -0.0678 0.3357 1.0000
3.000 0.6355 0.00852 0.00218 -0.0666 0.3063 1.0000
3.250 0.6577 0.00879 0.00234 -0.0654 0.2792 1.0000
3.500 0.6799 0.00906 0.00250 -0.0643 0.2493 1.0000
3.750 0.7006 0.00946 0.00271 -0.0630 0.2034 1.0000
4.000 0.7209 0.00989 0.00296 -0.0616 0.1638 1.0000
4.250 0.7419 0.01028 0.00324 -0.0603 0.1352 1.0000
4.500 0.7636 0.01062 0.00350 -0.0591 0.1131 1.0000
4.750 0.7847 0.01101 0.00378 -0.0578 0.0878 1.0000
5.000 0.8059 0.01139 0.00408 -0.0566 0.0674 1.0000
5.250 0.8277 0.01172 0.00438 -0.0554 0.0572 1.0000
5.500 0.8496 0.01204 0.00469 -0.0543 0.0514 1.0000
5.750 0.8713 0.01239 0.00506 -0.0531 0.0470 1.0000
6.000 0.8927 0.01276 0.00543 -0.0519 0.0433 1.0000
6.250 0.9146 0.01308 0.00579 -0.0508 0.0402 1.0000
6.500 0.9375 0.01330 0.00607 -0.0499 0.0365 1.0000
6.750 0.9589 0.01365 0.00642 -0.0487 0.0313 1.0000
7.000 0.9798 0.01404 0.00675 -0.0476 0.0194 1.0000
7.250 0.9978 0.01470 0.00735 -0.0458 0.0098 1.0000
7.500 1.0172 0.01524 0.00794 -0.0443 0.0077 1.0000
7.750 1.0362 0.01581 0.00858 -0.0427 0.0065 1.0000
8.000 1.0542 0.01645 0.00930 -0.0409 0.0057 1.0000
8.250 1.0703 0.01730 0.01026 -0.0388 0.0049 1.0000
8.500 1.0885 0.01789 0.01095 -0.0372 0.0047 1.0000
8.750 1.1062 0.01853 0.01166 -0.0355 0.0043 1.0000
9.000 1.1245 0.01906 0.01225 -0.0340 0.0038 1.0000
9.250 1.1401 0.01983 0.01312 -0.0320 0.0035 1.0000
9.500 1.1555 0.02059 0.01399 -0.0301 0.0033 1.0000
9.750 1.1688 0.02150 0.01500 -0.0278 0.0032 1.0000
10.000 1.1806 0.02247 0.01608 -0.0254 0.0030 1.0000
10.250 1.1859 0.02373 0.01747 -0.0219 0.0029 1.0000
10.500 1.1894 0.02511 0.01899 -0.0182 0.0028 1.0000
10.750 1.1986 0.02610 0.02010 -0.0156 0.0027 1.0000
11.000 1.2068 0.02719 0.02132 -0.0130 0.0026 1.0000
11.250 1.2133 0.02845 0.02273 -0.0104 0.0025 1.0000
11.500 1.2172 0.02997 0.02441 -0.0078 0.0025 1.0000
11.750 1.2209 0.03154 0.02614 -0.0054 0.0024 1.0000
12.000 1.2216 0.03344 0.02820 -0.0030 0.0025 1.0000
12.250 1.2222 0.03543 0.03037 -0.0009 0.0024 1.0000
12.500 1.2203 0.03773 0.03285 0.0009 0.0024 1.0000
12.750 1.2183 0.04020 0.03554 0.0022 0.0022 1.0000
13.000 1.2116 0.04332 0.03885 0.0032 0.0023 1.0000
13.250 1.2050 0.04665 0.04238 0.0036 0.0022 1.0000
13.500 1.1971 0.05037 0.04628 0.0034 0.0021 1.0000
13.750 1.1836 0.05512 0.05124 0.0025 0.0022 1.0000
14.000 1.1716 0.06003 0.05633 0.0009 0.0022 1.0000
14.250 1.1551 0.06610 0.06260 -0.0017 0.0021 1.0000
14.500 1.1403 0.07237 0.06903 -0.0049 0.0021 1.0000
14.750 1.1222 0.07982 0.07666 -0.0090 0.0021 1.0000
15.000 1.1030 0.08805 0.08505 -0.0137 0.0022 1.0000
15.250 1.0796 0.09761 0.09476 -0.0190 0.0022 1.0000
15.500 1.0572 0.10727 0.10455 -0.0242 0.0022 1.0000
15.750 1.0337 0.11745 0.11481 -0.0294 0.0023 1.0000
16.000 1.0093 0.12828 0.12574 -0.0347 0.0023 1.0000
16.250 0.9793 0.14101 0.13856 -0.0408 0.0024 1.0000
16.500 0.9470 0.15533 0.15297 -0.0475 0.0025 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 795 AIRFOIL (goe795-il)