GOE 456 AIRFOIL (goe456-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: GOE 456 AIRFOIL (goe456-il) Reynolds number: 500,000 Max Cl/Cd: 115.61 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe456-il-500000-n5.txt Download as CSV file: xf-goe456-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 456 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3002 0.12037 0.11809 -0.0341 1.0000 0.0046
-10.000 -0.2983 0.11745 0.11521 -0.0341 1.0000 0.0043
-9.750 -0.2969 0.11466 0.11244 -0.0340 1.0000 0.0042
-9.500 -0.2898 0.11136 0.10915 -0.0355 0.9993 0.0042
-9.250 -0.2760 0.10722 0.10501 -0.0391 0.9970 0.0041
-9.000 -0.2629 0.10313 0.10093 -0.0425 0.9930 0.0044
-8.750 -0.2494 0.09938 0.09717 -0.0459 0.9888 0.0043
-8.000 -0.1997 0.08987 0.08768 -0.0586 0.9707 0.0060
-7.750 -0.1825 0.08574 0.08355 -0.0632 0.9661 0.0060
-7.500 -0.1649 0.08177 0.07957 -0.0678 0.9589 0.0060
-7.250 -0.1443 0.07742 0.07521 -0.0735 0.9524 0.0060
-7.000 -0.1237 0.07313 0.07090 -0.0793 0.9433 0.0060
-6.750 -0.1075 0.06775 0.06549 -0.0847 0.9333 0.0056
-6.500 -0.0882 0.06324 0.06094 -0.0907 0.9216 0.0054
-6.250 -0.0702 0.05907 0.05672 -0.0960 0.9092 0.0052
-6.000 -0.0521 0.05483 0.05241 -0.1011 0.8969 0.0051
-5.750 -0.0328 0.05062 0.04811 -0.1060 0.8859 0.0050
-5.500 -0.0117 0.04603 0.04342 -0.1110 0.8756 0.0049
-5.250 0.0117 0.04135 0.03862 -0.1158 0.8661 0.0047
-5.000 0.0378 0.03604 0.03314 -0.1206 0.8574 0.0046
-4.750 0.0660 0.02900 0.02581 -0.1252 0.8492 0.0045
-4.500 0.0922 0.02192 0.01815 -0.1279 0.8413 0.0044
-4.250 0.1172 0.01767 0.01323 -0.1284 0.8341 0.0044
-4.000 0.1425 0.01458 0.00949 -0.1280 0.8267 0.0048
-3.750 0.1689 0.01297 0.00743 -0.1275 0.8196 0.0057
-3.500 0.1948 0.01197 0.00623 -0.1273 0.8116 0.0066
-3.250 0.2213 0.01175 0.00596 -0.1272 0.8037 0.0079
-3.000 0.2480 0.01167 0.00578 -0.1271 0.7953 0.0109
-2.750 0.2747 0.01080 0.00470 -0.1267 0.7870 0.0123
-2.500 0.3015 0.01040 0.00410 -0.1263 0.7783 0.0140
-2.250 0.3279 0.01000 0.00358 -0.1260 0.7683 0.0162
-2.000 0.3540 0.00952 0.00303 -0.1257 0.7579 0.0180
-1.750 0.3804 0.00925 0.00268 -0.1254 0.7472 0.0197
-1.500 0.4068 0.00906 0.00241 -0.1251 0.7362 0.0216
-1.250 0.4332 0.00885 0.00209 -0.1247 0.7243 0.0212
-0.750 0.4860 0.00862 0.00156 -0.1240 0.6985 0.0203
-0.500 0.5124 0.00859 0.00141 -0.1237 0.6860 0.0200
-0.250 0.5388 0.00857 0.00131 -0.1233 0.6743 0.0197
0.000 0.5649 0.00859 0.00124 -0.1229 0.6609 0.0195
0.250 0.5906 0.00865 0.00119 -0.1225 0.6455 0.0194
0.500 0.6163 0.00872 0.00118 -0.1221 0.6308 0.0197
0.750 0.6423 0.00879 0.00117 -0.1217 0.6183 0.0202
1.000 0.6683 0.00885 0.00118 -0.1214 0.6066 0.0219
1.250 0.6940 0.00868 0.00124 -0.1211 0.5963 0.1371
1.500 0.7196 0.00863 0.00136 -0.1208 0.5864 0.1997
1.750 0.7439 0.00819 0.00161 -0.1205 0.5765 0.4599
2.000 0.7765 0.00725 0.00176 -0.1218 0.5648 1.0000
2.250 0.8019 0.00739 0.00184 -0.1214 0.5521 1.0000
2.500 0.8275 0.00753 0.00193 -0.1210 0.5412 1.0000
2.750 0.8532 0.00767 0.00204 -0.1206 0.5322 1.0000
3.000 0.8792 0.00778 0.00215 -0.1203 0.5233 1.0000
3.250 0.9049 0.00792 0.00233 -0.1200 0.5146 1.0000
3.750 0.9549 0.00826 0.00263 -0.1190 0.4847 1.0000
4.000 0.9672 0.00930 0.00300 -0.1163 0.3567 1.0000
4.250 0.9805 0.01049 0.00363 -0.1141 0.2491 1.0000
4.500 0.9869 0.01235 0.00456 -0.1108 0.0879 1.0000
4.750 1.0011 0.01352 0.00539 -0.1085 0.0046 1.0000
5.000 1.0226 0.01405 0.00606 -0.1073 0.0028 1.0000
5.250 1.0441 0.01454 0.00665 -0.1062 0.0027 1.0000
5.500 1.0640 0.01518 0.00739 -0.1047 0.0025 1.0000
5.750 1.0821 0.01596 0.00827 -0.1030 0.0024 1.0000
6.000 1.0986 0.01685 0.00926 -0.1010 0.0023 1.0000
6.250 1.1134 0.01781 0.01031 -0.0987 0.0023 1.0000
6.500 1.1258 0.01892 0.01164 -0.0960 0.0023 1.0000
6.750 1.1370 0.02011 0.01292 -0.0932 0.0023 1.0000
7.000 1.1466 0.02136 0.01426 -0.0901 0.0023 1.0000
7.250 1.1576 0.02272 0.01569 -0.0873 0.0024 1.0000
7.500 1.1719 0.02424 0.01730 -0.0851 0.0024 1.0000
7.750 1.1915 0.02604 0.01919 -0.0839 0.0024 1.0000
8.000 1.2176 0.02828 0.02156 -0.0838 0.0025 1.0000
8.250 1.2440 0.03087 0.02431 -0.0837 0.0026 1.0000
8.500 1.2681 0.03435 0.02801 -0.0833 0.0028 1.0000
9.500 1.3327 0.04235 0.03678 -0.0761 0.0036 1.0000
9.750 1.3382 0.04512 0.03982 -0.0732 0.0037 1.0000
10.000 1.3391 0.04774 0.04267 -0.0697 0.0037 1.0000
10.250 1.3348 0.05024 0.04539 -0.0657 0.0037 1.0000
10.500 1.3273 0.05290 0.04835 -0.0617 0.0037 1.0000
10.750 1.3175 0.05564 0.05130 -0.0581 0.0037 1.0000
11.000 1.3062 0.05851 0.05436 -0.0549 0.0036 1.0000
11.250 1.2930 0.06176 0.05781 -0.0523 0.0036 1.0000
11.500 1.2791 0.06521 0.06145 -0.0503 0.0036 1.0000
11.750 1.2641 0.06905 0.06547 -0.0489 0.0036 1.0000
12.000 1.2481 0.07327 0.06987 -0.0482 0.0036 1.0000
12.250 1.2312 0.07798 0.07476 -0.0482 0.0036 1.0000
12.500 1.2137 0.08312 0.08007 -0.0491 0.0036 1.0000
12.750 1.1957 0.08876 0.08588 -0.0507 0.0037 1.0000
13.000 1.1773 0.09494 0.09221 -0.0531 0.0037 1.0000
13.250 1.1583 0.10180 0.09923 -0.0564 0.0037 1.0000
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Polar data table (+)
Polar graphs
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