GOE 363 AIRFOIL (goe363-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 363 AIRFOIL (goe363-il) Reynolds number: 100,000 Max Cl/Cd: 53.33 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe363-il-100000.txt Download as CSV file: xf-goe363-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 363 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
------ -------- --------- --------- -------- -------- --------
-7.000 -0.3499 0.10237 0.09853 -0.0116 1.0000 0.0946
-6.750 -0.3670 0.10148 0.09772 -0.0113 1.0000 0.0964
-6.500 -0.3841 0.10082 0.09711 -0.0180 1.0000 0.0980
-6.250 -0.3669 0.09520 0.09153 -0.0190 0.9963 0.0997
-6.000 -0.3440 0.09133 0.08764 -0.0183 0.9922 0.1030
-5.750 -0.2993 0.08649 0.08258 -0.0433 0.9805 0.1124
-5.500 -0.2867 0.08174 0.07794 -0.0384 0.9759 0.1140
-5.250 -0.2597 0.07826 0.07445 -0.0393 0.9710 0.1194
-5.000 -0.2250 0.07331 0.06935 -0.0521 0.9607 0.1285
-4.750 -0.1956 0.06971 0.06575 -0.0534 0.9560 0.1332
-4.500 -0.1577 0.06519 0.06099 -0.0646 0.9453 0.1434
-4.250 -0.1246 0.06154 0.05735 -0.0665 0.9409 0.1482
-4.000 -0.0918 0.05785 0.05347 -0.0728 0.9301 0.1599
-3.750 -0.0451 0.05429 0.04962 -0.0805 0.9233 0.1740
-3.500 0.0150 0.04312 0.03735 -0.0919 0.9167 0.1129
-3.250 0.0679 0.03649 0.02984 -0.0980 0.9128 0.0980
-3.000 0.1062 0.03349 0.02649 -0.1006 0.9051 0.0965
-2.750 0.1529 0.03089 0.02322 -0.1039 0.8986 0.0981
-2.500 0.1943 0.02909 0.02125 -0.1063 0.8893 0.1017
-2.250 0.2524 0.02701 0.01885 -0.1106 0.8816 0.1068
-2.000 0.2988 0.02542 0.01705 -0.1127 0.8692 0.1160
-1.750 0.3426 0.02397 0.01549 -0.1141 0.8584 0.1295
-1.500 0.3864 0.02248 0.01406 -0.1156 0.8510 0.1616
-1.250 0.4164 0.02168 0.01340 -0.1154 0.8394 0.2218
-1.000 0.4513 0.02104 0.01281 -0.1157 0.8304 0.2568
-0.750 0.4837 0.02050 0.01232 -0.1157 0.8201 0.2937
-0.500 0.5116 0.02000 0.01190 -0.1149 0.8073 0.3251
-0.250 0.5406 0.01936 0.01134 -0.1142 0.7952 0.3537
0.000 0.5727 0.01852 0.01063 -0.1138 0.7845 0.3961
0.250 0.6053 0.01650 0.00999 -0.1130 0.7722 1.0000
0.500 0.6329 0.01638 0.00957 -0.1121 0.7574 1.0000
0.750 0.6602 0.01624 0.00922 -0.1111 0.7421 1.0000
1.000 0.6874 0.01609 0.00888 -0.1102 0.7261 1.0000
1.250 0.7149 0.01591 0.00852 -0.1093 0.7092 1.0000
1.500 0.7395 0.01586 0.00833 -0.1081 0.6885 1.0000
1.750 0.7659 0.01574 0.00804 -0.1070 0.6673 1.0000
2.000 0.7919 0.01569 0.00779 -0.1060 0.6437 1.0000
2.250 0.8180 0.01571 0.00757 -0.1050 0.6182 1.0000
2.500 0.8425 0.01590 0.00751 -0.1038 0.5891 1.0000
2.750 0.8666 0.01625 0.00756 -0.1026 0.5585 1.0000
3.000 0.8900 0.01675 0.00777 -0.1015 0.5279 1.0000
3.250 0.9131 0.01733 0.00808 -0.1004 0.4993 1.0000
3.500 0.9366 0.01790 0.00839 -0.0995 0.4753 1.0000
3.750 0.9594 0.01841 0.00875 -0.0986 0.4535 1.0000
4.000 0.9827 0.01889 0.00909 -0.0977 0.4356 1.0000
4.250 1.0064 0.01938 0.00944 -0.0970 0.4204 1.0000
4.500 1.0306 0.01989 0.00983 -0.0964 0.4077 1.0000
4.750 1.0554 0.02042 0.01025 -0.0959 0.3973 1.0000
5.000 1.0802 0.02095 0.01070 -0.0954 0.3874 1.0000
5.250 1.1043 0.02151 0.01122 -0.0949 0.3777 1.0000
5.500 1.1298 0.02208 0.01164 -0.0946 0.3693 1.0000
5.750 1.1532 0.02266 0.01229 -0.0939 0.3607 1.0000
6.000 1.1794 0.02331 0.01280 -0.0938 0.3538 1.0000
6.250 1.2026 0.02394 0.01354 -0.0932 0.3464 1.0000
6.500 1.2291 0.02462 0.01411 -0.0931 0.3400 1.0000
6.750 1.2513 0.02532 0.01495 -0.0924 0.3330 1.0000
7.000 1.2765 0.02600 0.01558 -0.0921 0.3262 1.0000
7.250 1.2993 0.02678 0.01646 -0.0915 0.3195 1.0000
7.500 1.3224 0.02749 0.01721 -0.0909 0.3123 1.0000
7.750 1.3459 0.02834 0.01807 -0.0905 0.3055 1.0000
8.000 1.3666 0.02911 0.01896 -0.0896 0.2980 1.0000
8.250 1.3913 0.03003 0.01985 -0.0893 0.2913 1.0000
8.500 1.4091 0.03092 0.02095 -0.0880 0.2837 1.0000
8.750 1.4348 0.03186 0.02178 -0.0880 0.2767 1.0000
9.000 1.4482 0.03264 0.02281 -0.0860 0.2680 1.0000
9.250 1.4686 0.03333 0.02347 -0.0851 0.2597 1.0000
9.500 1.4829 0.03391 0.02418 -0.0834 0.2513 1.0000
9.750 1.5000 0.03463 0.02494 -0.0821 0.2438 1.0000
10.000 1.5130 0.03519 0.02562 -0.0802 0.2363 1.0000
10.250 1.5318 0.03590 0.02633 -0.0792 0.2303 1.0000
10.500 1.5407 0.03686 0.02757 -0.0768 0.2244 1.0000
10.750 1.5587 0.03728 0.02795 -0.0756 0.2189 1.0000
11.000 1.5632 0.03808 0.02896 -0.0727 0.2128 1.0000
11.250 1.5674 0.03858 0.02959 -0.0697 0.2068 1.0000
11.500 1.5810 0.03901 0.02998 -0.0680 0.2021 1.0000
11.750 1.5781 0.04028 0.03159 -0.0645 0.1978 1.0000
12.000 1.5807 0.04122 0.03270 -0.0618 0.1931 1.0000
12.250 1.5865 0.04172 0.03317 -0.0596 0.1882 1.0000
12.500 1.5783 0.04342 0.03521 -0.0566 0.1828 1.0000
12.750 1.5770 0.04472 0.03666 -0.0545 0.1774 1.0000
13.000 1.5767 0.04624 0.03831 -0.0528 0.1728 1.0000
13.250 1.5747 0.04831 0.04067 -0.0513 0.1678 1.0000
13.500 1.5712 0.05030 0.04278 -0.0500 0.1620 1.0000
13.750 1.5677 0.05285 0.04558 -0.0491 0.1565 1.0000
14.000 1.5624 0.05561 0.04849 -0.0485 0.1504 1.0000
14.250 1.5543 0.05898 0.05202 -0.0482 0.1434 1.0000
14.500 1.5431 0.06290 0.05602 -0.0484 0.1369 1.0000
14.750 1.5298 0.06756 0.06083 -0.0490 0.1301 1.0000
15.000 1.5136 0.07284 0.06620 -0.0501 0.1247 1.0000
15.250 1.4968 0.07864 0.07217 -0.0516 0.1200 1.0000
15.500 1.4777 0.08505 0.07878 -0.0536 0.1160 1.0000
15.750 1.4557 0.09209 0.08596 -0.0560 0.1129 1.0000
16.000 1.4328 0.09949 0.09345 -0.0588 0.1102 1.0000
16.250 1.4100 0.10723 0.10144 -0.0618 0.1081 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 363 AIRFOIL (goe363-il)