GOE 403 AIRFOIL (goe403-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 403 AIRFOIL (goe403-il) Reynolds number: 1,000,000 Max Cl/Cd: 116.08 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe403-il-1000000.txt Download as CSV file: xf-goe403-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 403 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3661 0.10737 0.10574 -0.0209 1.0000 0.0076
-9.250 -0.3621 0.10394 0.10232 -0.0219 1.0000 0.0076
-9.000 -0.3657 0.09886 0.09727 -0.0231 1.0000 0.0079
-8.750 -0.3592 0.09681 0.09523 -0.0232 1.0000 0.0084
-8.500 -0.3559 0.09380 0.09224 -0.0240 1.0000 0.0084
-8.250 -0.3508 0.09158 0.09003 -0.0243 1.0000 0.0089
-8.000 -0.3476 0.08914 0.08762 -0.0247 1.0000 0.0097
-7.750 -0.3473 0.08652 0.08502 -0.0248 1.0000 0.0100
-6.250 -0.2417 0.01564 0.01242 -0.0942 0.9613 0.0079
-6.000 -0.2129 0.01184 0.00788 -0.0956 0.9499 0.0089
-5.750 -0.1814 0.01126 0.00716 -0.0964 0.9357 0.0098
-5.500 -0.1522 0.01068 0.00639 -0.0967 0.9170 0.0108
-5.250 -0.1252 0.01013 0.00563 -0.0965 0.8966 0.0116
-5.000 -0.0989 0.00977 0.00508 -0.0960 0.8784 0.0121
-4.750 -0.0734 0.00897 0.00409 -0.0955 0.8639 0.0151
-4.500 -0.0462 0.00908 0.00417 -0.0952 0.8518 0.0172
-4.250 -0.0188 0.00928 0.00433 -0.0950 0.8416 0.0184
-3.750 0.0346 0.00874 0.00362 -0.0944 0.8234 0.0253
-3.500 0.0621 0.00890 0.00373 -0.0943 0.8153 0.0269
-3.250 0.0887 0.00857 0.00329 -0.0940 0.8070 0.0287
-3.000 0.1150 0.00801 0.00262 -0.0936 0.7990 0.0312
-2.500 0.1690 0.00766 0.00217 -0.0931 0.7824 0.0358
-2.250 0.1961 0.00754 0.00199 -0.0929 0.7739 0.0379
-2.000 0.2231 0.00743 0.00181 -0.0926 0.7643 0.0388
-1.750 0.2503 0.00731 0.00164 -0.0924 0.7541 0.0393
-1.500 0.2771 0.00706 0.00130 -0.0921 0.7429 0.0410
-1.250 0.3040 0.00692 0.00109 -0.0918 0.7298 0.0434
-1.000 0.3308 0.00683 0.00095 -0.0914 0.7142 0.0478
-0.750 0.3571 0.00670 0.00085 -0.0911 0.6952 0.0747
-0.500 0.3835 0.00667 0.00082 -0.0907 0.6718 0.1047
-0.250 0.4096 0.00672 0.00080 -0.0903 0.6450 0.1184
0.000 0.4356 0.00679 0.00078 -0.0899 0.6164 0.1318
0.250 0.4615 0.00685 0.00079 -0.0895 0.5899 0.1511
0.500 0.4876 0.00687 0.00082 -0.0892 0.5688 0.1908
0.750 0.5135 0.00676 0.00089 -0.0890 0.5517 0.2865
1.000 0.5378 0.00634 0.00100 -0.0886 0.5372 0.5224
1.250 0.5754 0.00539 0.00109 -0.0909 0.5223 1.0000
1.500 0.6016 0.00551 0.00115 -0.0905 0.5104 1.0000
1.750 0.6275 0.00567 0.00120 -0.0901 0.4916 1.0000
2.000 0.6533 0.00583 0.00126 -0.0897 0.4725 1.0000
2.250 0.6796 0.00596 0.00132 -0.0894 0.4595 1.0000
2.500 0.7058 0.00609 0.00140 -0.0890 0.4421 1.0000
2.750 0.7313 0.00630 0.00148 -0.0886 0.4153 1.0000
3.000 0.7560 0.00660 0.00159 -0.0880 0.3757 1.0000
3.250 0.7777 0.00723 0.00182 -0.0871 0.2941 1.0000
3.500 0.7867 0.00949 0.00282 -0.0843 0.0238 1.0000
3.750 0.8125 0.00971 0.00307 -0.0838 0.0160 1.0000
4.000 0.8382 0.00996 0.00336 -0.0834 0.0145 1.0000
4.250 0.8633 0.01026 0.00371 -0.0828 0.0131 1.0000
4.500 0.8877 0.01069 0.00420 -0.0821 0.0115 1.0000
4.750 0.9105 0.01132 0.00492 -0.0811 0.0106 1.0000
5.000 0.9320 0.01210 0.00582 -0.0799 0.0097 1.0000
5.250 0.9557 0.01254 0.00630 -0.0791 0.0094 1.0000
5.500 0.9786 0.01308 0.00688 -0.0782 0.0089 1.0000
5.750 1.0006 0.01370 0.00756 -0.0772 0.0084 1.0000
6.000 1.0212 0.01446 0.00838 -0.0759 0.0078 1.0000
6.250 1.0404 0.01540 0.00937 -0.0743 0.0074 1.0000
6.500 1.0568 0.01683 0.01091 -0.0721 0.0076 1.0000
7.000 1.0754 0.01359 0.00800 -0.0669 0.0098 1.0000
7.250 1.0969 0.01526 0.00981 -0.0659 0.0098 1.0000
7.500 1.1169 0.01699 0.01170 -0.0647 0.0098 1.0000
7.750 1.1357 0.01830 0.01317 -0.0634 0.0096 1.0000
8.000 1.1569 0.01684 0.01177 -0.0621 0.0083 1.0000
8.250 1.1757 0.01813 0.01320 -0.0608 0.0075 1.0000
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Polar data table (+)
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