GOE 437 AIRFOIL (goe437-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 437 AIRFOIL (goe437-il) Reynolds number: 500,000 Max Cl/Cd: 105.75 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe437-il-500000-n5.txt Download as CSV file: xf-goe437-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 437 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2323 0.10356 0.10147 -0.0303 1.0000 0.0079
-9.750 -0.2306 0.10065 0.09858 -0.0304 1.0000 0.0081
-9.500 -0.2293 0.09775 0.09571 -0.0303 1.0000 0.0085
-9.250 -0.2258 0.09460 0.09257 -0.0311 0.9995 0.0088
-9.000 -0.2174 0.09084 0.08881 -0.0338 0.9976 0.0091
-8.750 -0.2085 0.08713 0.08511 -0.0363 0.9948 0.0093
-8.500 -0.2003 0.08325 0.08123 -0.0386 0.9915 0.0094
-8.250 -0.1908 0.07924 0.07723 -0.0412 0.9882 0.0094
-8.000 -0.1824 0.07527 0.07326 -0.0436 0.9839 0.0094
-7.750 -0.1730 0.07115 0.06914 -0.0463 0.9795 0.0095
-7.250 -0.2504 0.08346 0.08137 -0.0420 0.9832 0.0094
-6.250 -0.1656 0.06231 0.06016 -0.0693 0.9555 0.0057
-6.000 -0.1385 0.05758 0.05539 -0.0764 0.9450 0.0054
-5.750 -0.1086 0.05232 0.05005 -0.0842 0.9325 0.0053
-5.500 -0.0789 0.04700 0.04462 -0.0911 0.9170 0.0052
-5.250 -0.0490 0.04100 0.03844 -0.0975 0.8989 0.0053
-5.000 -0.0203 0.03469 0.03186 -0.1025 0.8785 0.0053
-4.750 0.0061 0.02123 0.01761 -0.1072 0.8600 0.0055
-4.500 0.0304 0.01657 0.01217 -0.1075 0.8430 0.0060
-4.250 0.0543 0.01397 0.00897 -0.1071 0.8265 0.0063
-4.000 0.0797 0.01266 0.00729 -0.1066 0.8115 0.0066
-3.750 0.1054 0.01179 0.00612 -0.1061 0.7962 0.0070
-3.500 0.1311 0.01111 0.00518 -0.1055 0.7792 0.0076
-3.250 0.1566 0.01054 0.00436 -0.1049 0.7618 0.0084
-3.000 0.1825 0.01013 0.00374 -0.1043 0.7462 0.0092
-2.750 0.2085 0.00955 0.00296 -0.1038 0.7341 0.0116
-2.500 0.2351 0.00922 0.00252 -0.1034 0.7238 0.0142
-2.250 0.2617 0.00903 0.00221 -0.1030 0.7133 0.0178
-2.000 0.2882 0.00885 0.00199 -0.1027 0.7026 0.0303
-1.750 0.3149 0.00874 0.00184 -0.1024 0.6913 0.0403
-1.500 0.3415 0.00864 0.00168 -0.1021 0.6786 0.0494
-1.250 0.3679 0.00854 0.00155 -0.1018 0.6644 0.0620
-1.000 0.3943 0.00848 0.00148 -0.1014 0.6487 0.0813
-0.750 0.4205 0.00849 0.00142 -0.1010 0.6291 0.0983
-0.500 0.4463 0.00854 0.00134 -0.1006 0.6032 0.1076
-0.250 0.4718 0.00862 0.00129 -0.1001 0.5773 0.1155
0.000 0.4974 0.00870 0.00127 -0.0996 0.5572 0.1251
0.250 0.5234 0.00877 0.00127 -0.0993 0.5423 0.1350
0.750 0.5757 0.00885 0.00133 -0.0987 0.5186 0.1739
1.000 0.6018 0.00881 0.00139 -0.0984 0.5082 0.2277
1.250 0.6226 0.00793 0.00157 -0.0975 0.4993 0.6588
1.750 0.6900 0.00749 0.00168 -0.1000 0.4813 1.0000
2.250 0.7422 0.00774 0.00183 -0.0994 0.4633 1.0000
2.500 0.7683 0.00788 0.00192 -0.0990 0.4537 1.0000
3.000 0.8203 0.00814 0.00213 -0.0983 0.4343 1.0000
3.250 0.8462 0.00829 0.00225 -0.0980 0.4246 1.0000
3.500 0.8719 0.00845 0.00238 -0.0976 0.4138 1.0000
3.750 0.8976 0.00861 0.00254 -0.0972 0.4024 1.0000
4.000 0.9232 0.00878 0.00269 -0.0969 0.3917 1.0000
4.250 0.9486 0.00897 0.00287 -0.0964 0.3808 1.0000
4.500 0.9729 0.00924 0.00307 -0.0959 0.3606 1.0000
4.750 0.9962 0.00961 0.00333 -0.0952 0.3325 1.0000
5.000 1.0194 0.01000 0.00360 -0.0944 0.3040 1.0000
5.250 1.0409 0.01055 0.00393 -0.0935 0.2580 1.0000
5.500 1.0550 0.01190 0.00465 -0.0916 0.1442 1.0000
5.750 1.0645 0.01383 0.00590 -0.0889 0.0163 1.0000
6.000 1.0872 0.01429 0.00643 -0.0880 0.0115 1.0000
6.250 1.1092 0.01483 0.00707 -0.0870 0.0088 1.0000
6.500 1.1314 0.01530 0.00761 -0.0861 0.0074 1.0000
6.750 1.1523 0.01589 0.00828 -0.0850 0.0065 1.0000
7.000 1.1696 0.01684 0.00935 -0.0833 0.0058 1.0000
7.250 1.1890 0.01753 0.01013 -0.0820 0.0053 1.0000
7.500 1.2078 0.01821 0.01092 -0.0806 0.0047 1.0000
7.750 1.2262 0.01887 0.01164 -0.0792 0.0042 1.0000
8.000 1.2408 0.01983 0.01267 -0.0773 0.0039 1.0000
8.250 1.2495 0.02119 0.01415 -0.0744 0.0037 1.0000
8.500 1.2612 0.02224 0.01532 -0.0719 0.0035 1.0000
8.750 1.2676 0.02345 0.01665 -0.0686 0.0033 1.0000
9.000 1.2722 0.02480 0.01812 -0.0651 0.0032 1.0000
9.250 1.2770 0.02628 0.01972 -0.0618 0.0031 1.0000
9.500 1.2823 0.02791 0.02147 -0.0589 0.0030 1.0000
9.750 1.2886 0.02971 0.02341 -0.0562 0.0029 1.0000
10.000 1.2963 0.03169 0.02553 -0.0539 0.0028 1.0000
10.250 1.3047 0.03362 0.02761 -0.0518 0.0027 1.0000
10.500 1.3118 0.03492 0.02902 -0.0500 0.0025 1.0000
10.750 1.3172 0.03636 0.03056 -0.0483 0.0024 1.0000
11.000 1.3209 0.03837 0.03276 -0.0465 0.0023 1.0000
11.250 1.3211 0.04110 0.03566 -0.0447 0.0022 1.0000
11.500 1.3233 0.04392 0.03875 -0.0430 0.0021 1.0000
11.750 1.3220 0.04738 0.04247 -0.0413 0.0020 1.0000
12.000 1.3158 0.05193 0.04736 -0.0395 0.0019 1.0000
12.250 1.3073 0.05624 0.05192 -0.0385 0.0019 1.0000
12.500 1.2969 0.06058 0.05648 -0.0381 0.0019 1.0000
12.750 1.2848 0.06517 0.06128 -0.0383 0.0019 1.0000
13.000 1.2704 0.07044 0.06676 -0.0390 0.0019 1.0000
13.250 1.2539 0.07631 0.07285 -0.0405 0.0018 1.0000
13.500 1.2386 0.08215 0.07886 -0.0427 0.0019 1.0000
13.750 1.2223 0.08863 0.08552 -0.0455 0.0019 1.0000
14.000 1.2039 0.09608 0.09315 -0.0491 0.0018 1.0000
14.250 1.1891 0.10299 0.10021 -0.0529 0.0019 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 437 AIRFOIL (goe437-il)