GOE 444 AIRFOIL (goe444-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 444 AIRFOIL (goe444-il) Reynolds number: 1,000,000 Max Cl/Cd: 42.27 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe444-il-1000000.txt Download as CSV file: xf-goe444-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6379 0.08310 0.08151 -0.0025 1.0000 0.0059
-8.250 -0.6440 0.07810 0.07654 -0.0071 1.0000 0.0057
-8.000 -0.6524 0.07317 0.07159 -0.0110 1.0000 0.0058
-7.750 -0.6541 0.06811 0.06647 -0.0138 1.0000 0.0057
-7.500 -0.6529 0.06325 0.06154 -0.0155 1.0000 0.0058
-7.250 -0.6492 0.05858 0.05676 -0.0164 1.0000 0.0060
-7.000 -0.6429 0.05419 0.05225 -0.0165 1.0000 0.0062
-6.750 -0.6335 0.05019 0.04811 -0.0159 1.0000 0.0065
-6.500 -0.6158 0.04758 0.04538 -0.0149 1.0000 0.0072
-6.250 -0.6014 0.04465 0.04230 -0.0137 1.0000 0.0074
-6.000 -0.5891 0.04135 0.03882 -0.0122 1.0000 0.0075
-5.750 -0.5765 0.03799 0.03527 -0.0105 1.0000 0.0075
-5.500 -0.5632 0.03478 0.03184 -0.0085 1.0000 0.0076
-5.250 -0.5490 0.03171 0.02854 -0.0064 1.0000 0.0076
-5.000 -0.5343 0.02871 0.02530 -0.0042 1.0000 0.0076
-4.750 -0.5187 0.02584 0.02217 -0.0019 1.0000 0.0076
-4.500 -0.5025 0.02305 0.01911 0.0005 1.0000 0.0076
-4.250 -0.4944 0.01549 0.01080 0.0052 1.0000 0.0054
-4.000 -0.4733 0.01324 0.00829 0.0073 1.0000 0.0058
-3.750 -0.4501 0.01285 0.00783 0.0084 1.0000 0.0071
-3.500 -0.4294 0.01132 0.00610 0.0100 1.0000 0.0086
-3.250 -0.4085 0.01043 0.00518 0.0113 1.0000 0.0104
-3.000 -0.3881 0.00941 0.00405 0.0131 1.0000 0.0098
-2.750 -0.3673 0.00871 0.00326 0.0147 1.0000 0.0093
-2.500 -0.3408 0.00806 0.00250 0.0151 0.9993 0.0097
-2.250 -0.3051 0.00757 0.00191 0.0134 0.9971 0.0135
-2.000 -0.2682 0.00735 0.00164 0.0113 0.9948 0.0162
-1.750 -0.2385 0.00627 0.00123 0.0104 0.9918 0.2180
-1.500 -0.2081 0.00552 0.00108 0.0093 0.9880 0.3945
-1.250 -0.1789 0.00458 0.00099 0.0084 0.9842 0.6337
-1.000 -0.1460 0.00415 0.00095 0.0071 0.9812 0.7466
-0.750 -0.1218 0.00368 0.00094 0.0082 0.9741 0.8636
-0.500 -0.0861 0.00353 0.00091 0.0066 0.9704 0.9043
-0.250 -0.0470 0.00348 0.00092 0.0042 0.9667 0.9316
0.000 0.0000 0.00349 0.00095 0.0000 0.9570 0.9569
0.250 0.0467 0.00349 0.00092 -0.0041 0.9322 0.9671
0.500 0.0862 0.00354 0.00091 -0.0066 0.9038 0.9704
0.750 0.1214 0.00370 0.00093 -0.0081 0.8576 0.9740
1.000 0.1466 0.00413 0.00096 -0.0073 0.7515 0.9811
1.250 0.1784 0.00465 0.00098 -0.0083 0.6135 0.9843
1.500 0.2080 0.00552 0.00108 -0.0093 0.3950 0.9880
1.750 0.2386 0.00626 0.00122 -0.0104 0.2223 0.9918
2.000 0.2684 0.00734 0.00163 -0.0114 0.0162 0.9948
2.250 0.3054 0.00757 0.00191 -0.0135 0.0133 0.9971
2.500 0.3408 0.00810 0.00254 -0.0151 0.0095 0.9993
2.750 0.3673 0.00869 0.00324 -0.0148 0.0094 1.0000
3.000 0.3880 0.00942 0.00406 -0.0131 0.0098 1.0000
3.250 0.4081 0.01061 0.00537 -0.0111 0.0108 1.0000
3.500 0.4292 0.01137 0.00614 -0.0100 0.0086 1.0000
3.750 0.4498 0.01298 0.00798 -0.0083 0.0072 1.0000
4.000 0.4731 0.01332 0.00838 -0.0072 0.0059 1.0000
4.250 0.4943 0.01543 0.01074 -0.0052 0.0055 1.0000
6.000 0.5892 0.04133 0.03880 0.0122 0.0075 1.0000
6.250 0.6014 0.04468 0.04233 0.0137 0.0074 1.0000
6.500 0.6152 0.04771 0.04552 0.0149 0.0072 1.0000
6.750 0.6336 0.05017 0.04809 0.0159 0.0065 1.0000
7.000 0.6428 0.05422 0.05228 0.0164 0.0062 1.0000
7.250 0.6492 0.05859 0.05678 0.0163 0.0060 1.0000
7.500 0.6532 0.06325 0.06154 0.0155 0.0058 1.0000
7.750 0.6541 0.06815 0.06652 0.0137 0.0058 1.0000
8.000 0.6531 0.07322 0.07165 0.0109 0.0057 1.0000
8.250 0.6441 0.07826 0.07670 0.0069 0.0057 1.0000
8.500 0.6386 0.08322 0.08163 0.0023 0.0058 1.0000
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Polar data table (+)
Polar graphs
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