GOE 602 MOD. AIRFOIL (goe602m-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 602 MOD. AIRFOIL (goe602m-il) Reynolds number: 1,000,000 Max Cl/Cd: 108.89 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe602m-il-1000000-n5.txt Download as CSV file: xf-goe602m-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 MOD. 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.4101 0.12550 0.12379 -0.0294 1.0000 0.0034
-12.000 -0.4046 0.12303 0.12133 -0.0301 1.0000 0.0036
-11.500 -0.7518 0.02909 0.02663 -0.0984 0.9874 0.0035
-11.250 -0.7342 0.02563 0.02287 -0.1014 0.9837 0.0035
-11.000 -0.7126 0.02328 0.02029 -0.1032 0.9803 0.0037
-10.750 -0.6869 0.02144 0.01825 -0.1050 0.9782 0.0038
-10.500 -0.6588 0.01994 0.01656 -0.1067 0.9767 0.0039
-10.250 -0.6353 0.01879 0.01526 -0.1070 0.9722 0.0040
-10.000 -0.6077 0.01766 0.01395 -0.1080 0.9692 0.0042
-9.750 -0.5785 0.01661 0.01274 -0.1093 0.9668 0.0044
-9.500 -0.5501 0.01568 0.01166 -0.1101 0.9639 0.0045
-9.250 -0.5244 0.01487 0.01072 -0.1103 0.9589 0.0047
-9.000 -0.4952 0.01414 0.00985 -0.1111 0.9553 0.0049
-8.750 -0.4654 0.01345 0.00903 -0.1120 0.9520 0.0051
-8.500 -0.4400 0.01269 0.00817 -0.1119 0.9461 0.0055
-8.250 -0.4115 0.01213 0.00752 -0.1124 0.9413 0.0059
-8.000 -0.3832 0.01166 0.00697 -0.1127 0.9367 0.0064
-7.750 -0.3565 0.01123 0.00643 -0.1127 0.9307 0.0068
-7.500 -0.3285 0.01082 0.00591 -0.1128 0.9253 0.0072
-7.250 -0.3021 0.01040 0.00540 -0.1127 0.9196 0.0078
-7.000 -0.2753 0.01000 0.00494 -0.1125 0.9136 0.0087
-6.750 -0.2482 0.00969 0.00457 -0.1125 0.9080 0.0097
-6.500 -0.2217 0.00941 0.00422 -0.1122 0.9016 0.0106
-6.000 -0.1678 0.00897 0.00375 -0.1119 0.8888 0.0152
-5.750 -0.1407 0.00879 0.00353 -0.1118 0.8816 0.0172
-5.500 -0.1137 0.00869 0.00343 -0.1116 0.8734 0.0196
-5.250 -0.0864 0.00862 0.00332 -0.1114 0.8655 0.0219
-5.000 -0.0592 0.00856 0.00322 -0.1112 0.8569 0.0234
-4.750 -0.0319 0.00851 0.00313 -0.1110 0.8490 0.0242
-4.500 -0.0047 0.00849 0.00306 -0.1108 0.8401 0.0246
-4.250 0.0217 0.00824 0.00274 -0.1105 0.8307 0.0257
-4.000 0.0479 0.00799 0.00240 -0.1102 0.8208 0.0271
-3.750 0.0743 0.00783 0.00216 -0.1098 0.8092 0.0283
-3.250 0.1271 0.00760 0.00177 -0.1091 0.7828 0.0294
-3.000 0.1534 0.00752 0.00161 -0.1087 0.7684 0.0298
-2.750 0.1797 0.00747 0.00148 -0.1083 0.7535 0.0307
-2.500 0.2058 0.00742 0.00135 -0.1079 0.7381 0.0313
-2.250 0.2321 0.00739 0.00124 -0.1075 0.7232 0.0320
-2.000 0.2584 0.00737 0.00115 -0.1071 0.7078 0.0333
-1.750 0.2847 0.00737 0.00108 -0.1067 0.6931 0.0348
-1.500 0.3110 0.00735 0.00101 -0.1063 0.6780 0.0424
-1.250 0.3367 0.00721 0.00096 -0.1059 0.6628 0.0875
-1.000 0.3628 0.00719 0.00094 -0.1055 0.6475 0.1059
-0.750 0.3889 0.00717 0.00091 -0.1052 0.6324 0.1248
-0.500 0.4138 0.00692 0.00089 -0.1048 0.6167 0.2250
-0.250 0.4395 0.00692 0.00092 -0.1043 0.5985 0.2660
0.000 0.4645 0.00702 0.00095 -0.1038 0.5714 0.2861
0.250 0.4888 0.00719 0.00100 -0.1030 0.5363 0.2977
0.500 0.5139 0.00733 0.00105 -0.1025 0.5085 0.3085
0.750 0.5395 0.00742 0.00110 -0.1021 0.4876 0.3257
1.000 0.5653 0.00747 0.00115 -0.1017 0.4714 0.3510
1.250 0.5903 0.00734 0.00124 -0.1013 0.4570 0.4459
1.500 0.6155 0.00732 0.00135 -0.1008 0.4426 0.5174
1.750 0.6413 0.00738 0.00143 -0.1004 0.4288 0.5512
2.000 0.6672 0.00748 0.00152 -0.1000 0.4147 0.5704
2.250 0.6930 0.00755 0.00161 -0.0997 0.4023 0.5941
2.500 0.7184 0.00761 0.00170 -0.0992 0.3885 0.6244
2.750 0.7425 0.00757 0.00182 -0.0985 0.3744 0.6923
3.000 0.7609 0.00718 0.00198 -0.0964 0.3621 0.9045
3.250 0.8069 0.00741 0.00215 -0.1007 0.3345 1.0000
3.500 0.8291 0.00782 0.00235 -0.0997 0.2918 1.0000
3.750 0.8521 0.00817 0.00255 -0.0988 0.2596 1.0000
4.000 0.8743 0.00859 0.00279 -0.0978 0.2225 1.0000
4.250 0.8965 0.00902 0.00304 -0.0969 0.1884 1.0000
4.500 0.9151 0.00975 0.00344 -0.0953 0.1240 1.0000
4.750 0.9352 0.01037 0.00385 -0.0940 0.0836 1.0000
5.000 0.9594 0.01063 0.00408 -0.0934 0.0739 1.0000
5.250 0.9777 0.01139 0.00458 -0.0918 0.0229 1.0000
5.500 1.0012 0.01171 0.00488 -0.0910 0.0150 1.0000
5.750 1.0247 0.01201 0.00520 -0.0902 0.0120 1.0000
6.000 1.0482 0.01230 0.00551 -0.0895 0.0103 1.0000
6.250 1.0716 0.01259 0.00582 -0.0887 0.0093 1.0000
6.500 1.0943 0.01292 0.00616 -0.0879 0.0083 1.0000
6.750 1.1160 0.01333 0.00660 -0.0868 0.0073 1.0000
7.000 1.1379 0.01371 0.00703 -0.0858 0.0068 1.0000
7.250 1.1600 0.01405 0.00739 -0.0849 0.0063 1.0000
7.500 1.1815 0.01442 0.00778 -0.0838 0.0058 1.0000
7.750 1.2025 0.01481 0.00819 -0.0827 0.0053 1.0000
8.000 1.2218 0.01532 0.00873 -0.0813 0.0049 1.0000
8.250 1.2403 0.01587 0.00933 -0.0798 0.0046 1.0000
8.500 1.2593 0.01635 0.00985 -0.0783 0.0044 1.0000
8.750 1.2769 0.01684 0.01042 -0.0767 0.0042 1.0000
9.000 1.2922 0.01739 0.01103 -0.0746 0.0040 1.0000
9.250 1.3080 0.01787 0.01154 -0.0726 0.0038 1.0000
9.500 1.3212 0.01850 0.01224 -0.0702 0.0037 1.0000
9.750 1.3351 0.01910 0.01288 -0.0681 0.0036 1.0000
10.000 1.3487 0.01972 0.01354 -0.0659 0.0034 1.0000
10.250 1.3604 0.02045 0.01433 -0.0636 0.0033 1.0000
10.500 1.3666 0.02153 0.01549 -0.0605 0.0030 1.0000
10.750 1.3781 0.02231 0.01634 -0.0583 0.0030 1.0000
11.000 1.3874 0.02324 0.01735 -0.0560 0.0029 1.0000
11.250 1.3951 0.02431 0.01851 -0.0535 0.0028 1.0000
11.500 1.4039 0.02533 0.01964 -0.0514 0.0027 1.0000
11.750 1.4109 0.02652 0.02091 -0.0492 0.0026 1.0000
12.000 1.4164 0.02787 0.02236 -0.0470 0.0026 1.0000
12.250 1.4206 0.02938 0.02396 -0.0448 0.0025 1.0000
12.500 1.4256 0.03089 0.02557 -0.0430 0.0024 1.0000
12.750 1.4294 0.03258 0.02735 -0.0412 0.0024 1.0000
13.000 1.4298 0.03465 0.02953 -0.0394 0.0024 1.0000
13.250 1.4287 0.03699 0.03199 -0.0379 0.0023 1.0000
13.500 1.4338 0.03881 0.03389 -0.0369 0.0023 1.0000
13.750 1.4329 0.04136 0.03655 -0.0359 0.0022 1.0000
14.000 1.4351 0.04369 0.03897 -0.0353 0.0022 1.0000
14.250 1.4354 0.04632 0.04170 -0.0348 0.0021 1.0000
14.500 1.4346 0.04920 0.04469 -0.0346 0.0021 1.0000
14.750 1.4325 0.05241 0.04802 -0.0347 0.0021 1.0000
15.000 1.4246 0.05648 0.05223 -0.0350 0.0021 1.0000
15.250 1.4151 0.06099 0.05689 -0.0356 0.0021 1.0000
15.500 1.4217 0.06347 0.05941 -0.0364 0.0020 1.0000
15.750 1.4101 0.06859 0.06468 -0.0376 0.0020 1.0000
16.000 1.4019 0.07342 0.06963 -0.0391 0.0020 1.0000
16.250 1.3957 0.07816 0.07448 -0.0408 0.0019 1.0000
16.500 1.3836 0.08397 0.08043 -0.0429 0.0019 1.0000
16.750 1.3708 0.09005 0.08665 -0.0453 0.0019 1.0000
17.000 1.3616 0.09578 0.09249 -0.0479 0.0019 1.0000
17.250 1.3481 0.10238 0.09923 -0.0508 0.0019 1.0000
17.500 1.3362 0.10890 0.10587 -0.0540 0.0019 1.0000
17.750 1.3227 0.11587 0.11298 -0.0575 0.0019 1.0000
18.000 1.3087 0.12315 0.12038 -0.0613 0.0019 1.0000
18.250 1.2960 0.13033 0.12770 -0.0651 0.0019 1.0000
18.500 1.2837 0.13757 0.13505 -0.0691 0.0019 1.0000
18.750 1.2696 0.14540 0.14301 -0.0736 0.0019 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 602 MOD. AIRFOIL (goe602m-il)