GOE 362 AIRFOIL (goe362-il) Xfoil prediction polar at RE=500,000 Ncrit=5
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Airfoil: GOE 362 AIRFOIL (goe362-il) Reynolds number: 500,000 Max Cl/Cd: 108.15 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe362-il-500000-n5.txt Download as CSV file: xf-goe362-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 362 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2875 0.09881 0.09659 -0.0258 1.0000 0.0091
-8.250 -0.2842 0.09623 0.09404 -0.0260 1.0000 0.0093
-8.000 -0.2840 0.09387 0.09172 -0.0267 1.0000 0.0102
-7.250 -0.2458 0.08372 0.08161 -0.0375 0.9815 0.0104
-7.000 -0.2225 0.07929 0.07717 -0.0438 0.9744 0.0104
-6.750 -0.1963 0.07471 0.07258 -0.0508 0.9673 0.0104
-6.250 -0.1493 0.06614 0.06397 -0.0616 0.9489 0.0093
-5.750 -0.0918 0.05356 0.05123 -0.0795 0.9145 0.0071
-5.500 -0.0659 0.04884 0.04640 -0.0856 0.8959 0.0070
-5.000 0.0117 0.01785 0.01388 -0.1130 0.8609 0.0072
-4.750 0.0378 0.01618 0.01186 -0.1133 0.8434 0.0078
-4.500 0.0642 0.01509 0.01048 -0.1133 0.8268 0.0083
-4.250 0.0909 0.01358 0.00856 -0.1132 0.8124 0.0087
-4.000 0.1177 0.01232 0.00691 -0.1130 0.7984 0.0092
-3.750 0.1446 0.01146 0.00574 -0.1127 0.7843 0.0098
-3.500 0.1715 0.01086 0.00491 -0.1123 0.7703 0.0103
-3.250 0.1983 0.01012 0.00395 -0.1120 0.7564 0.0114
-3.000 0.2252 0.00992 0.00366 -0.1116 0.7414 0.0130
-2.750 0.2522 0.00964 0.00324 -0.1112 0.7267 0.0146
-2.500 0.2792 0.00938 0.00284 -0.1109 0.7124 0.0159
-2.250 0.3062 0.00907 0.00243 -0.1105 0.6980 0.0200
-2.000 0.3332 0.00897 0.00222 -0.1101 0.6836 0.0241
-1.750 0.3602 0.00881 0.00198 -0.1098 0.6702 0.0300
-1.500 0.3872 0.00875 0.00183 -0.1095 0.6574 0.0359
-1.250 0.4142 0.00865 0.00167 -0.1091 0.6446 0.0441
-1.000 0.4411 0.00859 0.00158 -0.1088 0.6321 0.0572
-0.750 0.4682 0.00854 0.00154 -0.1085 0.6198 0.0763
-0.500 0.4953 0.00855 0.00149 -0.1083 0.6072 0.0903
-0.250 0.5224 0.00857 0.00146 -0.1080 0.5945 0.1001
0.000 0.5493 0.00858 0.00142 -0.1077 0.5811 0.1093
0.250 0.5762 0.00860 0.00140 -0.1074 0.5661 0.1203
0.500 0.6030 0.00862 0.00139 -0.1071 0.5500 0.1348
0.750 0.6295 0.00861 0.00145 -0.1068 0.5317 0.1833
1.000 0.6559 0.00865 0.00148 -0.1065 0.5125 0.2214
1.250 0.6822 0.00871 0.00153 -0.1061 0.4921 0.2543
1.500 0.7082 0.00870 0.00163 -0.1058 0.4740 0.3379
1.750 0.7278 0.00783 0.00182 -0.1044 0.4595 0.8076
2.250 0.7904 0.00791 0.00196 -0.1057 0.4337 1.0000
2.500 0.8169 0.00808 0.00206 -0.1054 0.4236 1.0000
2.750 0.8434 0.00825 0.00217 -0.1051 0.4141 1.0000
3.000 0.8700 0.00841 0.00229 -0.1048 0.4057 1.0000
3.250 0.8964 0.00858 0.00242 -0.1044 0.3974 1.0000
3.500 0.9230 0.00873 0.00255 -0.1042 0.3896 1.0000
4.000 0.9756 0.00908 0.00285 -0.1035 0.3713 1.0000
4.250 1.0016 0.00928 0.00301 -0.1032 0.3585 1.0000
4.500 1.0274 0.00950 0.00318 -0.1028 0.3433 1.0000
4.750 1.0529 0.00974 0.00338 -0.1024 0.3268 1.0000
5.000 1.0783 0.00999 0.00357 -0.1020 0.3107 1.0000
5.250 1.1037 0.01024 0.00379 -0.1015 0.2963 1.0000
5.500 1.1286 0.01055 0.00404 -0.1011 0.2781 1.0000
5.750 1.1533 0.01086 0.00431 -0.1006 0.2595 1.0000
6.000 1.1773 0.01126 0.00462 -0.1000 0.2375 1.0000
6.250 1.2009 0.01169 0.00497 -0.0994 0.2135 1.0000
6.500 1.2240 0.01217 0.00537 -0.0987 0.1902 1.0000
6.750 1.2466 0.01272 0.00581 -0.0980 0.1646 1.0000
7.000 1.2673 0.01346 0.00637 -0.0970 0.1272 1.0000
7.250 1.2860 0.01442 0.00708 -0.0958 0.0877 1.0000
7.500 1.3071 0.01507 0.00766 -0.0949 0.0722 1.0000
7.750 1.3292 0.01558 0.00819 -0.0941 0.0622 1.0000
8.000 1.3510 0.01611 0.00871 -0.0933 0.0537 1.0000
8.250 1.3720 0.01669 0.00927 -0.0923 0.0453 1.0000
8.500 1.3924 0.01731 0.00986 -0.0913 0.0358 1.0000
8.750 1.4099 0.01819 0.01064 -0.0899 0.0182 1.0000
9.000 1.4270 0.01908 0.01152 -0.0884 0.0111 1.0000
9.250 1.4433 0.01998 0.01252 -0.0868 0.0085 1.0000
9.500 1.4610 0.02070 0.01333 -0.0854 0.0074 1.0000
9.750 1.4769 0.02153 0.01425 -0.0837 0.0066 1.0000
10.000 1.4901 0.02251 0.01532 -0.0818 0.0060 1.0000
10.250 1.5026 0.02343 0.01635 -0.0796 0.0055 1.0000
10.500 1.5127 0.02431 0.01734 -0.0771 0.0051 1.0000
10.750 1.5212 0.02527 0.01839 -0.0745 0.0047 1.0000
11.000 1.5287 0.02632 0.01953 -0.0719 0.0044 1.0000
11.250 1.5336 0.02760 0.02090 -0.0693 0.0042 1.0000
11.500 1.5333 0.02934 0.02276 -0.0664 0.0039 1.0000
11.750 1.5375 0.03086 0.02440 -0.0643 0.0038 1.0000
12.000 1.5403 0.03259 0.02626 -0.0623 0.0036 1.0000
12.250 1.5409 0.03463 0.02848 -0.0606 0.0036 1.0000
12.500 1.5411 0.03687 0.03085 -0.0591 0.0034 1.0000
12.750 1.5402 0.03938 0.03349 -0.0580 0.0033 1.0000
13.000 1.5381 0.04219 0.03644 -0.0572 0.0032 1.0000
13.250 1.5348 0.04531 0.03970 -0.0568 0.0031 1.0000
13.500 1.5311 0.04868 0.04320 -0.0567 0.0030 1.0000
13.750 1.5260 0.05240 0.04706 -0.0569 0.0030 1.0000
14.000 1.5201 0.05642 0.05121 -0.0575 0.0029 1.0000
14.250 1.5132 0.06070 0.05563 -0.0583 0.0028 1.0000
14.500 1.5049 0.06539 0.06046 -0.0595 0.0028 1.0000
14.750 1.4951 0.07051 0.06573 -0.0610 0.0028 1.0000
15.000 1.4858 0.07578 0.07113 -0.0628 0.0027 1.0000
15.250 1.4748 0.08154 0.07701 -0.0649 0.0027 1.0000
15.500 1.4642 0.08740 0.08302 -0.0672 0.0027 1.0000
15.750 1.4522 0.09364 0.08939 -0.0698 0.0026 1.0000
16.000 1.4399 0.10011 0.09599 -0.0726 0.0026 1.0000
16.250 1.4287 0.10652 0.10254 -0.0755 0.0026 1.0000
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Polar data table (+)
Polar graphs
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