GOE 445 AIRFOIL (goe445-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 445 AIRFOIL (goe445-il) Reynolds number: 50,000 Max Cl/Cd: 18.44 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe445-il-50000.txt Download as CSV file: xf-goe445-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 445 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.250 -0.5891 0.08542 0.07885 0.0258 1.0000 0.4097
-6.250 -0.6005 0.04858 0.04239 -0.0081 1.0000 0.2295
-6.000 -0.6354 0.05203 0.04451 -0.0124 1.0000 0.1563
-5.750 -0.6199 0.04668 0.03792 -0.0126 1.0000 0.1242
-5.500 -0.5998 0.04202 0.03310 -0.0117 1.0000 0.1159
-5.250 -0.5786 0.03826 0.02866 -0.0102 1.0000 0.1061
-5.000 -0.5555 0.03533 0.02471 -0.0082 1.0000 0.0980
-4.750 -0.5306 0.03205 0.02112 -0.0070 1.0000 0.0961
-4.500 -0.5042 0.02932 0.01799 -0.0058 1.0000 0.0956
-4.250 -0.4758 0.02700 0.01527 -0.0045 1.0000 0.0972
-4.000 -0.4487 0.02478 0.01305 -0.0036 1.0000 0.1097
-3.750 -0.4211 0.02281 0.01106 -0.0024 1.0000 0.1245
-3.500 -0.2218 0.01868 0.00936 -0.0196 1.0000 1.0000
-3.250 -0.2093 0.01795 0.00837 -0.0189 1.0000 1.0000
-3.000 -0.1963 0.01732 0.00751 -0.0179 1.0000 1.0000
-2.750 -0.1825 0.01680 0.00676 -0.0168 1.0000 1.0000
-2.500 -0.1680 0.01636 0.00612 -0.0155 1.0000 1.0000
-2.250 -0.1528 0.01601 0.00550 -0.0142 1.0000 1.0000
-2.000 -0.1369 0.01572 0.00504 -0.0127 1.0000 1.0000
-1.750 -0.1206 0.01548 0.00466 -0.0112 1.0000 1.0000
-1.500 -0.1040 0.01529 0.00434 -0.0097 1.0000 1.0000
-1.250 -0.0870 0.01514 0.00408 -0.0081 1.0000 1.0000
-1.000 -0.0699 0.01502 0.00387 -0.0065 1.0000 1.0000
-0.750 -0.0526 0.01493 0.00369 -0.0049 1.0000 1.0000
-0.500 -0.0352 0.01486 0.00357 -0.0032 1.0000 1.0000
-0.250 -0.0177 0.01483 0.00350 -0.0016 1.0000 1.0000
0.000 0.0000 0.01481 0.00348 0.0000 1.0000 1.0000
0.250 0.0177 0.01483 0.00350 0.0016 1.0000 1.0000
0.500 0.0352 0.01486 0.00357 0.0032 1.0000 1.0000
0.750 0.0526 0.01492 0.00369 0.0049 1.0000 1.0000
1.000 0.0699 0.01501 0.00387 0.0065 1.0000 1.0000
1.250 0.0871 0.01513 0.00408 0.0081 1.0000 1.0000
1.500 0.1040 0.01529 0.00434 0.0097 1.0000 1.0000
1.750 0.1207 0.01548 0.00465 0.0112 1.0000 1.0000
2.000 0.1370 0.01572 0.00504 0.0127 1.0000 1.0000
2.250 0.1529 0.01600 0.00549 0.0141 1.0000 1.0000
2.500 0.1681 0.01636 0.00603 0.0155 1.0000 1.0000
2.750 0.1826 0.01679 0.00676 0.0168 1.0000 1.0000
3.000 0.1964 0.01732 0.00750 0.0179 1.0000 1.0000
3.250 0.2095 0.01794 0.00837 0.0189 1.0000 1.0000
3.500 0.2220 0.01868 0.00936 0.0196 1.0000 1.0000
3.750 0.4210 0.02283 0.01107 0.0024 0.1244 1.0000
4.000 0.4487 0.02478 0.01305 0.0036 0.1096 1.0000
4.250 0.4758 0.02699 0.01527 0.0045 0.0972 1.0000
4.500 0.5041 0.02931 0.01797 0.0057 0.0957 1.0000
4.750 0.5305 0.03205 0.02111 0.0070 0.0961 1.0000
5.000 0.5555 0.03529 0.02468 0.0082 0.0981 1.0000
5.250 0.5786 0.03826 0.02865 0.0102 0.1061 1.0000
5.500 0.5999 0.04202 0.03309 0.0117 0.1160 1.0000
5.750 0.6186 0.04652 0.03828 0.0127 0.1305 1.0000
6.000 0.6354 0.05206 0.04455 0.0124 0.1565 1.0000
7.750 0.5990 0.09232 0.08566 -0.0239 0.3717 1.0000
8.000 0.6161 0.09671 0.09004 -0.0226 0.3524 1.0000
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Polar data table (+)
Polar graphs
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