GOE 602 AIRFOIL (goe602-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: GOE 602 AIRFOIL (goe602-il) Reynolds number: 500,000 Max Cl/Cd: 102.17 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe602-il-500000.txt Download as CSV file: xf-goe602-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 602 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4034 0.08842 0.08631 -0.0338 1.0000 0.0214
-8.500 -0.4141 0.08626 0.08420 -0.0321 1.0000 0.0215
-8.250 -0.4227 0.08389 0.08187 -0.0313 0.9997 0.0220
-8.000 -0.4105 0.07878 0.07676 -0.0379 0.9963 0.0227
-7.750 -0.3992 0.07261 0.07059 -0.0471 0.9909 0.0237
-7.500 -0.3675 0.05933 0.05716 -0.0693 0.9847 0.0258
-7.250 -0.3826 0.03812 0.03528 -0.0822 0.9743 0.0188
-7.000 -0.3638 0.02940 0.02578 -0.0866 0.9712 0.0190
-6.750 -0.3475 0.02591 0.02175 -0.0860 0.9644 0.0194
-6.500 -0.3252 0.02095 0.01616 -0.0872 0.9611 0.0203
-6.250 -0.2920 0.01969 0.01478 -0.0890 0.9595 0.0213
-6.000 -0.2567 0.01870 0.01366 -0.0910 0.9583 0.0227
-5.750 -0.2194 0.01801 0.01280 -0.0930 0.9574 0.0250
-5.500 -0.1955 0.01773 0.01234 -0.0921 0.9510 0.0260
-5.250 -0.1655 0.01491 0.00926 -0.0932 0.9485 0.0281
-5.000 -0.1303 0.01420 0.00851 -0.0948 0.9463 0.0303
-4.750 -0.0941 0.01354 0.00776 -0.0966 0.9443 0.0328
-4.500 -0.0663 0.01302 0.00714 -0.0964 0.9388 0.0343
-4.250 -0.0374 0.01196 0.00598 -0.0966 0.9340 0.0361
-4.000 -0.0054 0.01111 0.00511 -0.0976 0.9305 0.0390
-3.750 0.0209 0.01067 0.00462 -0.0971 0.9242 0.0411
-3.500 0.0501 0.01026 0.00416 -0.0973 0.9185 0.0433
-3.250 0.0802 0.00992 0.00376 -0.0977 0.9134 0.0452
-3.000 0.1057 0.00955 0.00334 -0.0971 0.9055 0.0469
-2.750 0.1352 0.00914 0.00288 -0.0973 0.8998 0.0512
-2.500 0.1605 0.00892 0.00264 -0.0966 0.8911 0.0564
-2.250 0.1885 0.00861 0.00238 -0.0965 0.8840 0.0754
-2.000 0.2121 0.00798 0.00217 -0.0958 0.8748 0.1933
-1.750 0.2369 0.00755 0.00208 -0.0954 0.8657 0.3161
-1.500 0.2636 0.00726 0.00197 -0.0951 0.8571 0.3872
-1.250 0.2864 0.00685 0.00192 -0.0941 0.8460 0.5019
-1.000 0.3085 0.00644 0.00191 -0.0929 0.8350 0.6406
-0.750 0.3312 0.00615 0.00188 -0.0915 0.8235 0.7355
-0.500 0.3525 0.00585 0.00191 -0.0896 0.8116 0.8480
-0.250 0.3883 0.00576 0.00191 -0.0908 0.7993 0.9396
0.000 0.4454 0.00579 0.00187 -0.0969 0.7861 0.9790
0.250 0.4934 0.00585 0.00181 -0.1012 0.7683 0.9951
0.500 0.5269 0.00592 0.00176 -0.1025 0.7472 1.0000
0.750 0.5478 0.00603 0.00173 -0.1010 0.7252 1.0000
1.000 0.5692 0.00615 0.00173 -0.0995 0.7047 1.0000
1.250 0.5909 0.00628 0.00175 -0.0982 0.6861 1.0000
1.500 0.6128 0.00641 0.00179 -0.0969 0.6675 1.0000
1.750 0.6350 0.00655 0.00184 -0.0957 0.6490 1.0000
2.000 0.6570 0.00670 0.00190 -0.0944 0.6299 1.0000
2.250 0.6790 0.00687 0.00198 -0.0932 0.6113 1.0000
2.500 0.7011 0.00704 0.00206 -0.0920 0.5901 1.0000
2.750 0.7229 0.00724 0.00215 -0.0907 0.5691 1.0000
3.000 0.7456 0.00741 0.00226 -0.0896 0.5503 1.0000
3.250 0.7687 0.00759 0.00238 -0.0886 0.5332 1.0000
3.500 0.7918 0.00778 0.00251 -0.0877 0.5160 1.0000
3.750 0.8148 0.00798 0.00265 -0.0867 0.4988 1.0000
4.000 0.8378 0.00820 0.00282 -0.0857 0.4816 1.0000
4.250 0.8602 0.00843 0.00298 -0.0847 0.4581 1.0000
4.500 0.8814 0.00875 0.00316 -0.0834 0.4278 1.0000
4.750 0.9034 0.00903 0.00335 -0.0823 0.4029 1.0000
5.000 0.9247 0.00937 0.00357 -0.0811 0.3715 1.0000
5.250 0.9453 0.00977 0.00382 -0.0798 0.3371 1.0000
5.500 0.9652 0.01024 0.00410 -0.0785 0.2958 1.0000
5.750 0.9835 0.01084 0.00446 -0.0769 0.2478 1.0000
6.000 0.9991 0.01168 0.00495 -0.0749 0.1842 1.0000
6.250 1.0079 0.01307 0.00577 -0.0719 0.0906 1.0000
6.500 1.0196 0.01424 0.00658 -0.0693 0.0308 1.0000
6.750 1.0379 0.01489 0.00729 -0.0675 0.0247 1.0000
7.000 1.0580 0.01537 0.00785 -0.0661 0.0231 1.0000
7.250 1.0765 0.01594 0.00849 -0.0645 0.0211 1.0000
7.500 1.0923 0.01668 0.00930 -0.0625 0.0194 1.0000
7.750 1.1001 0.01785 0.01059 -0.0591 0.0180 1.0000
8.000 1.1129 0.01855 0.01136 -0.0565 0.0175 1.0000
8.250 1.1254 0.01925 0.01213 -0.0539 0.0170 1.0000
8.500 1.1366 0.02006 0.01301 -0.0511 0.0163 1.0000
8.750 1.1469 0.02094 0.01396 -0.0484 0.0158 1.0000
9.000 1.1574 0.02184 0.01493 -0.0459 0.0150 1.0000
9.250 1.1682 0.02273 0.01585 -0.0435 0.0142 1.0000
9.500 1.1752 0.02396 0.01714 -0.0407 0.0138 1.0000
9.750 1.1803 0.02555 0.01877 -0.0379 0.0133 1.0000
10.000 1.1889 0.02832 0.02157 -0.0357 0.0128 1.0000
10.250 1.2014 0.02923 0.02260 -0.0338 0.0126 1.0000
10.500 1.2144 0.03040 0.02387 -0.0321 0.0123 1.0000
10.750 1.2286 0.03186 0.02546 -0.0306 0.0121 1.0000
11.000 1.2441 0.03363 0.02735 -0.0293 0.0120 1.0000
11.250 1.2583 0.03555 0.02943 -0.0279 0.0118 1.0000
11.500 1.2692 0.03748 0.03152 -0.0263 0.0116 1.0000
11.750 1.2763 0.03924 0.03344 -0.0244 0.0112 1.0000
12.000 1.2814 0.04109 0.03544 -0.0225 0.0109 1.0000
12.250 1.2856 0.04353 0.03807 -0.0207 0.0108 1.0000
12.500 1.2866 0.04647 0.04124 -0.0188 0.0109 1.0000
12.750 1.2821 0.05038 0.04542 -0.0168 0.0115 1.0000
13.000 1.2738 0.05410 0.04938 -0.0151 0.0117 1.0000
13.250 1.2627 0.05802 0.05351 -0.0137 0.0119 1.0000
13.500 1.2489 0.06233 0.05803 -0.0128 0.0121 1.0000
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Polar data table (+)
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