GOE 444 AIRFOIL (goe444-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 444 AIRFOIL (goe444-il) Reynolds number: 500,000 Max Cl/Cd: 45.93 at α=1.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe444-il-500000.txt Download as CSV file: xf-goe444-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 444 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6406 0.08071 0.07851 -0.0074 1.0000 0.0087
-8.250 -0.6485 0.07538 0.07318 -0.0124 1.0000 0.0086
-8.000 -0.6530 0.07063 0.06839 -0.0145 1.0000 0.0086
-7.750 -0.6539 0.06613 0.06381 -0.0159 1.0000 0.0087
-7.500 -0.6529 0.06180 0.05939 -0.0166 1.0000 0.0088
-7.250 -0.6498 0.05771 0.05519 -0.0168 1.0000 0.0089
-7.000 -0.6445 0.05380 0.05115 -0.0166 1.0000 0.0091
-6.750 -0.6371 0.05003 0.04723 -0.0161 1.0000 0.0093
-6.500 -0.6282 0.04643 0.04347 -0.0153 1.0000 0.0097
-6.250 -0.6176 0.04295 0.03981 -0.0141 1.0000 0.0101
-6.000 -0.6053 0.03963 0.03627 -0.0127 1.0000 0.0106
-5.750 -0.5918 0.03641 0.03283 -0.0111 1.0000 0.0112
-5.500 -0.5767 0.03338 0.02954 -0.0093 1.0000 0.0120
-5.250 -0.5600 0.03060 0.02649 -0.0073 1.0000 0.0131
-5.000 -0.5359 0.02967 0.02534 -0.0055 1.0000 0.0158
-4.750 -0.5142 0.02880 0.02420 -0.0036 1.0000 0.0165
-4.500 -0.4965 0.02629 0.02139 -0.0015 1.0000 0.0165
-4.000 -0.4398 0.00823 0.00294 0.0005 1.0000 0.0165
-3.750 -0.4410 0.01606 0.01004 0.0051 1.0000 0.0177
-3.250 -0.3953 0.01305 0.00680 0.0080 1.0000 0.0154
-3.000 -0.3742 0.01163 0.00529 0.0097 1.0000 0.0150
-2.750 -0.3538 0.01060 0.00417 0.0114 1.0000 0.0151
-2.500 -0.3326 0.00990 0.00339 0.0130 1.0000 0.0161
-2.250 -0.3121 0.00908 0.00243 0.0146 1.0000 0.0200
-2.000 -0.2897 0.00875 0.00206 0.0157 1.0000 0.0271
-1.750 -0.2768 0.00690 0.00155 0.0180 1.0000 0.3958
-1.500 -0.2636 0.00595 0.00151 0.0207 1.0000 0.6330
-1.250 -0.2516 0.00539 0.00160 0.0242 1.0000 0.7955
-1.000 -0.1676 0.00522 0.00186 0.0130 1.0000 0.9679
-0.750 -0.1221 0.00532 0.00190 0.0091 1.0000 0.9817
-0.500 -0.0699 0.00541 0.00194 0.0037 1.0000 0.9925
-0.250 -0.0194 0.00545 0.00195 -0.0014 1.0000 0.9991
0.000 0.0000 0.00544 0.00193 0.0000 1.0000 1.0000
0.250 0.0197 0.00545 0.00195 0.0013 0.9991 1.0000
0.500 0.0698 0.00541 0.00194 -0.0037 0.9925 1.0000
0.750 0.1215 0.00532 0.00190 -0.0089 0.9819 1.0000
1.000 0.1672 0.00522 0.00186 -0.0129 0.9682 1.0000
1.250 0.2503 0.00545 0.00157 -0.0239 0.7766 1.0000
1.500 0.2637 0.00594 0.00151 -0.0207 0.6340 1.0000
2.000 0.2897 0.00874 0.00205 -0.0158 0.0269 1.0000
2.250 0.3120 0.00909 0.00244 -0.0146 0.0197 1.0000
2.500 0.3327 0.00988 0.00337 -0.0130 0.0160 1.0000
2.750 0.3537 0.01059 0.00416 -0.0114 0.0152 1.0000
3.000 0.3742 0.01163 0.00529 -0.0097 0.0150 1.0000
3.250 0.3953 0.01306 0.00681 -0.0080 0.0154 1.0000
5.000 0.5065 0.01758 0.01352 0.0078 0.0163 1.0000
5.250 0.5331 0.01727 0.01341 0.0095 0.0137 1.0000
5.500 0.5486 0.01991 0.01633 0.0114 0.0125 1.0000
5.750 0.5622 0.02305 0.01972 0.0132 0.0116 1.0000
6.000 0.5742 0.02650 0.02340 0.0146 0.0110 1.0000
6.250 0.5845 0.03029 0.02740 0.0159 0.0104 1.0000
6.500 0.5930 0.03433 0.03163 0.0168 0.0101 1.0000
6.750 0.5990 0.03868 0.03614 0.0175 0.0098 1.0000
7.000 0.6028 0.04316 0.04076 0.0178 0.0095 1.0000
7.250 0.6038 0.04782 0.04554 0.0177 0.0094 1.0000
7.500 0.6009 0.05272 0.05054 0.0172 0.0093 1.0000
7.750 0.5941 0.05786 0.05575 0.0158 0.0093 1.0000
8.000 0.5766 0.06295 0.06089 0.0134 0.0095 1.0000
8.250 0.5619 0.06881 0.06672 0.0086 0.0096 1.0000
8.500 0.5545 0.07399 0.07188 0.0060 0.0096 1.0000
8.750 0.5492 0.07904 0.07692 0.0038 0.0096 1.0000
9.000 0.5454 0.08391 0.08177 0.0020 0.0096 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 444 AIRFOIL (goe444-il)