GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 457 AIRFOIL (goe457-il) Reynolds number: 500,000 Max Cl/Cd: 93.48 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe457-il-500000-n5.txt Download as CSV file: xf-goe457-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 457 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3680 0.09678 0.09454 -0.0236 1.0000 0.0078
-8.750 -0.3714 0.09372 0.09151 -0.0234 1.0000 0.0078
-8.500 -0.3770 0.09090 0.08873 -0.0227 1.0000 0.0079
-7.500 -0.3658 0.01978 0.01584 -0.0942 0.9676 0.0116
-7.250 -0.3330 0.01990 0.01601 -0.0953 0.9662 0.0122
-7.000 -0.3014 0.02026 0.01642 -0.0959 0.9641 0.0127
-6.750 -0.2738 0.02043 0.01659 -0.0958 0.9597 0.0134
-6.500 -0.2479 0.01641 0.01182 -0.0970 0.9559 0.0154
-6.250 -0.2141 0.01600 0.01134 -0.0983 0.9536 0.0162
-6.000 -0.1846 0.01602 0.01137 -0.0986 0.9493 0.0169
-5.750 -0.1550 0.01589 0.01119 -0.0990 0.9443 0.0179
-5.500 -0.1232 0.01509 0.01018 -0.0999 0.9408 0.0197
-5.250 -0.0915 0.01448 0.00936 -0.1006 0.9369 0.0209
-5.000 -0.0643 0.01354 0.00825 -0.1007 0.9303 0.0220
-4.750 -0.0316 0.01319 0.00786 -0.1017 0.9251 0.0231
-4.500 -0.0025 0.01281 0.00741 -0.1019 0.9185 0.0242
-4.250 0.0273 0.01235 0.00683 -0.1023 0.9121 0.0255
-4.000 0.0566 0.01192 0.00628 -0.1024 0.9054 0.0267
-3.750 0.0849 0.01145 0.00569 -0.1024 0.8974 0.0276
-3.500 0.1129 0.01107 0.00521 -0.1022 0.8883 0.0283
-3.250 0.1414 0.01084 0.00489 -0.1021 0.8782 0.0290
-3.000 0.1679 0.01035 0.00428 -0.1017 0.8657 0.0295
-2.750 0.1940 0.00978 0.00360 -0.1011 0.8499 0.0302
-2.500 0.2198 0.00940 0.00312 -0.1005 0.8254 0.0308
-2.250 0.2447 0.00920 0.00271 -0.0995 0.7772 0.0314
-2.000 0.2665 0.00924 0.00243 -0.0979 0.7169 0.0321
-1.750 0.2884 0.00931 0.00225 -0.0965 0.6657 0.0330
-1.500 0.3107 0.00941 0.00210 -0.0952 0.6152 0.0337
-1.250 0.3337 0.00949 0.00197 -0.0940 0.5744 0.0344
-1.000 0.3581 0.00954 0.00186 -0.0932 0.5477 0.0353
-0.750 0.3832 0.00956 0.00177 -0.0925 0.5282 0.0365
-0.500 0.4089 0.00958 0.00170 -0.0919 0.5134 0.0377
-0.250 0.4348 0.00960 0.00164 -0.0914 0.5005 0.0392
0.000 0.4609 0.00962 0.00161 -0.0909 0.4894 0.0429
0.500 0.5131 0.00959 0.00163 -0.0900 0.4686 0.0794
0.750 0.5392 0.00960 0.00165 -0.0896 0.4597 0.0965
1.000 0.5643 0.00938 0.00172 -0.0891 0.4501 0.2140
1.250 0.5900 0.00930 0.00179 -0.0886 0.4407 0.2868
1.500 0.6149 0.00916 0.00187 -0.0881 0.4311 0.3865
2.000 0.6850 0.00808 0.00212 -0.0914 0.4102 1.0000
2.250 0.7103 0.00823 0.00218 -0.0908 0.3982 1.0000
2.500 0.7355 0.00839 0.00227 -0.0902 0.3856 1.0000
2.750 0.7608 0.00855 0.00235 -0.0896 0.3734 1.0000
3.000 0.7859 0.00872 0.00245 -0.0890 0.3594 1.0000
3.250 0.8106 0.00893 0.00256 -0.0883 0.3419 1.0000
3.500 0.8353 0.00915 0.00269 -0.0876 0.3226 1.0000
3.750 0.8603 0.00934 0.00283 -0.0870 0.3099 1.0000
4.000 0.8853 0.00955 0.00297 -0.0864 0.2988 1.0000
4.250 0.9101 0.00976 0.00314 -0.0858 0.2853 1.0000
4.750 0.9591 0.01026 0.00351 -0.0845 0.2550 1.0000
5.000 0.9834 0.01053 0.00372 -0.0838 0.2390 1.0000
5.250 1.0076 0.01081 0.00395 -0.0832 0.2233 1.0000
5.500 1.0316 0.01111 0.00419 -0.0825 0.2070 1.0000
5.750 1.0543 0.01154 0.00450 -0.0816 0.1792 1.0000
6.250 1.0954 0.01285 0.00539 -0.0793 0.1070 1.0000
6.500 1.1180 0.01328 0.00577 -0.0784 0.0947 1.0000
6.750 1.1401 0.01374 0.00618 -0.0775 0.0815 1.0000
7.000 1.1624 0.01416 0.00656 -0.0766 0.0719 1.0000
7.250 1.1847 0.01457 0.00694 -0.0757 0.0637 1.0000
7.500 1.2068 0.01500 0.00733 -0.0748 0.0554 1.0000
7.750 1.2279 0.01549 0.00778 -0.0737 0.0456 1.0000
8.000 1.2435 0.01649 0.00859 -0.0719 0.0177 1.0000
8.250 1.2635 0.01706 0.00921 -0.0706 0.0132 1.0000
8.500 1.2834 0.01760 0.00983 -0.0693 0.0116 1.0000
8.750 1.3017 0.01826 0.01058 -0.0677 0.0102 1.0000
9.000 1.3198 0.01890 0.01132 -0.0662 0.0094 1.0000
9.250 1.3384 0.01946 0.01197 -0.0647 0.0088 1.0000
9.500 1.3559 0.02008 0.01269 -0.0631 0.0082 1.0000
9.750 1.3724 0.02073 0.01342 -0.0614 0.0076 1.0000
10.000 1.3859 0.02150 0.01427 -0.0592 0.0070 1.0000
10.250 1.3947 0.02248 0.01536 -0.0563 0.0066 1.0000
10.500 1.4070 0.02320 0.01616 -0.0539 0.0064 1.0000
10.750 1.4167 0.02408 0.01714 -0.0513 0.0062 1.0000
11.000 1.4260 0.02499 0.01815 -0.0487 0.0059 1.0000
11.250 1.4333 0.02603 0.01931 -0.0460 0.0057 1.0000
11.500 1.4392 0.02719 0.02058 -0.0434 0.0055 1.0000
11.750 1.4439 0.02847 0.02196 -0.0408 0.0054 1.0000
12.000 1.4486 0.02980 0.02339 -0.0384 0.0051 1.0000
12.250 1.4507 0.03139 0.02508 -0.0361 0.0050 1.0000
12.500 1.4525 0.03309 0.02688 -0.0340 0.0049 1.0000
12.750 1.4501 0.03526 0.02916 -0.0320 0.0047 1.0000
13.000 1.4451 0.03785 0.03187 -0.0304 0.0046 1.0000
13.250 1.4374 0.04100 0.03516 -0.0293 0.0045 1.0000
13.500 1.4356 0.04381 0.03810 -0.0289 0.0044 1.0000
13.750 1.4288 0.04751 0.04194 -0.0291 0.0044 1.0000
14.000 1.4245 0.05119 0.04576 -0.0297 0.0043 1.0000
14.250 1.4190 0.05530 0.05002 -0.0309 0.0042 1.0000
14.500 1.4100 0.06017 0.05503 -0.0324 0.0041 1.0000
14.750 1.3993 0.06538 0.06039 -0.0343 0.0041 1.0000
15.000 1.3886 0.07070 0.06586 -0.0363 0.0040 1.0000
15.250 1.3773 0.07623 0.07152 -0.0385 0.0040 1.0000
15.500 1.3646 0.08204 0.07746 -0.0408 0.0040 1.0000
15.750 1.3536 0.08777 0.08331 -0.0432 0.0039 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 457 AIRFOIL (goe457-il)