GOE 602 AIRFOIL (goe602-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 602 AIRFOIL (goe602-il) Reynolds number: 200,000 Max Cl/Cd: 73.96 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe602-il-200000-n5.txt Download as CSV file: xf-goe602-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3952 0.08772 0.08440 -0.0374 1.0000 0.0156
-8.500 -0.4074 0.08465 0.08141 -0.0365 1.0000 0.0155
-8.250 -0.4240 0.08180 0.07863 -0.0350 1.0000 0.0153
-8.000 -0.4243 0.07615 0.07303 -0.0407 0.9952 0.0153
-7.750 -0.4141 0.06660 0.06347 -0.0540 0.9871 0.0150
-7.500 -0.4075 0.05168 0.04831 -0.0706 0.9775 0.0147
-7.250 -0.4014 0.04075 0.03685 -0.0776 0.9688 0.0148
-7.000 -0.3835 0.03447 0.03003 -0.0810 0.9638 0.0153
-6.750 -0.3643 0.03079 0.02588 -0.0819 0.9576 0.0165
-6.500 -0.3378 0.02710 0.02155 -0.0835 0.9542 0.0185
-6.250 -0.3133 0.02430 0.01812 -0.0837 0.9495 0.0199
-6.000 -0.2890 0.02204 0.01553 -0.0839 0.9446 0.0213
-5.750 -0.2573 0.02129 0.01464 -0.0852 0.9416 0.0234
-5.500 -0.2235 0.02024 0.01334 -0.0867 0.9394 0.0265
-5.250 -0.1969 0.01920 0.01201 -0.0864 0.9341 0.0288
-5.000 -0.1678 0.01795 0.01057 -0.0870 0.9298 0.0316
-4.750 -0.1345 0.01709 0.00962 -0.0883 0.9269 0.0347
-4.500 -0.0996 0.01639 0.00879 -0.0898 0.9244 0.0383
-4.250 -0.0749 0.01583 0.00811 -0.0891 0.9172 0.0403
-4.000 -0.0431 0.01505 0.00724 -0.0899 0.9129 0.0422
-3.750 -0.0107 0.01429 0.00645 -0.0910 0.9089 0.0458
-3.500 0.0151 0.01383 0.00593 -0.0905 0.9011 0.0479
-3.250 0.0486 0.01335 0.00536 -0.0916 0.8964 0.0506
-3.000 0.0752 0.01303 0.00494 -0.0913 0.8882 0.0532
-2.750 0.1071 0.01261 0.00449 -0.0920 0.8823 0.0585
-2.500 0.1338 0.01232 0.00420 -0.0917 0.8738 0.0688
-2.250 0.1647 0.01186 0.00388 -0.0924 0.8674 0.1120
-2.000 0.1896 0.01142 0.00366 -0.0919 0.8578 0.1754
-1.750 0.2171 0.01094 0.00352 -0.0920 0.8495 0.2859
-1.500 0.2444 0.01060 0.00336 -0.0920 0.8404 0.3579
-1.250 0.2691 0.01021 0.00328 -0.0914 0.8299 0.4505
-1.000 0.2929 0.00975 0.00324 -0.0904 0.8197 0.5768
-0.750 0.3166 0.00935 0.00319 -0.0892 0.8092 0.7019
-0.500 0.3468 0.00894 0.00323 -0.0887 0.7982 0.8724
-0.250 0.4128 0.00881 0.00308 -0.0964 0.7875 0.9812
0.000 0.4526 0.00882 0.00296 -0.0990 0.7744 1.0000
0.250 0.4775 0.00885 0.00287 -0.0983 0.7596 1.0000
0.500 0.5023 0.00891 0.00281 -0.0976 0.7445 1.0000
0.750 0.5269 0.00899 0.00278 -0.0968 0.7290 1.0000
1.000 0.5514 0.00909 0.00276 -0.0961 0.7131 1.0000
1.250 0.5758 0.00921 0.00277 -0.0952 0.6966 1.0000
1.500 0.5999 0.00934 0.00279 -0.0944 0.6800 1.0000
1.750 0.6233 0.00949 0.00283 -0.0934 0.6603 1.0000
2.000 0.6466 0.00966 0.00289 -0.0924 0.6398 1.0000
2.250 0.6694 0.00985 0.00296 -0.0913 0.6170 1.0000
2.500 0.6923 0.01005 0.00305 -0.0902 0.5964 1.0000
2.750 0.7153 0.01026 0.00318 -0.0892 0.5770 1.0000
3.000 0.7386 0.01046 0.00331 -0.0883 0.5592 1.0000
3.250 0.7620 0.01067 0.00346 -0.0874 0.5423 1.0000
3.500 0.7852 0.01089 0.00363 -0.0865 0.5253 1.0000
3.750 0.8082 0.01113 0.00383 -0.0855 0.5084 1.0000
4.250 0.8533 0.01165 0.00425 -0.0835 0.4708 1.0000
4.500 0.8760 0.01192 0.00450 -0.0826 0.4535 1.0000
4.750 0.8988 0.01219 0.00476 -0.0816 0.4377 1.0000
5.000 0.9216 0.01246 0.00504 -0.0807 0.4227 1.0000
5.250 0.9435 0.01278 0.00533 -0.0796 0.4044 1.0000
5.500 0.9624 0.01321 0.00565 -0.0780 0.3702 1.0000
5.750 0.9787 0.01378 0.00600 -0.0761 0.3215 1.0000
6.000 0.9950 0.01442 0.00642 -0.0742 0.2750 1.0000
6.250 1.0087 0.01529 0.00696 -0.0720 0.2174 1.0000
6.500 1.0227 0.01621 0.00759 -0.0699 0.1611 1.0000
6.750 1.0320 0.01752 0.00848 -0.0672 0.0961 1.0000
7.250 1.0551 0.01976 0.01034 -0.0622 0.0267 1.0000
7.500 1.0709 0.02045 0.01117 -0.0602 0.0228 1.0000
7.750 1.0838 0.02127 0.01210 -0.0578 0.0199 1.0000
8.000 1.0945 0.02226 0.01323 -0.0551 0.0181 1.0000
8.250 1.1032 0.02334 0.01446 -0.0522 0.0169 1.0000
8.500 1.1153 0.02419 0.01543 -0.0499 0.0160 1.0000
8.750 1.1257 0.02516 0.01656 -0.0475 0.0149 1.0000
9.000 1.1344 0.02626 0.01777 -0.0450 0.0140 1.0000
9.250 1.1409 0.02753 0.01915 -0.0423 0.0133 1.0000
9.500 1.1468 0.02891 0.02062 -0.0398 0.0129 1.0000
9.750 1.1516 0.03043 0.02223 -0.0373 0.0124 1.0000
10.000 1.1559 0.03209 0.02397 -0.0350 0.0120 1.0000
10.250 1.1590 0.03402 0.02596 -0.0327 0.0115 1.0000
10.500 1.1620 0.03643 0.02842 -0.0306 0.0110 1.0000
10.750 1.1715 0.03794 0.03006 -0.0291 0.0105 1.0000
11.000 1.1806 0.03948 0.03174 -0.0277 0.0100 1.0000
11.250 1.1899 0.04137 0.03375 -0.0263 0.0097 1.0000
11.500 1.1991 0.04337 0.03593 -0.0249 0.0094 1.0000
11.750 1.2071 0.04550 0.03822 -0.0237 0.0092 1.0000
12.000 1.2139 0.04782 0.04072 -0.0224 0.0090 1.0000
12.250 1.2188 0.05029 0.04337 -0.0213 0.0088 1.0000
12.500 1.2214 0.05298 0.04625 -0.0203 0.0086 1.0000
12.750 1.2220 0.05589 0.04936 -0.0194 0.0085 1.0000
13.000 1.2203 0.05905 0.05273 -0.0187 0.0084 1.0000
13.250 1.2164 0.06251 0.05640 -0.0183 0.0083 1.0000
13.500 1.2110 0.06616 0.06025 -0.0183 0.0083 1.0000
13.750 1.2047 0.06998 0.06424 -0.0186 0.0081 1.0000
14.000 1.1959 0.07429 0.06876 -0.0193 0.0081 1.0000
14.250 1.1855 0.07897 0.07364 -0.0205 0.0080 1.0000
14.500 1.1772 0.08334 0.07814 -0.0221 0.0078 1.0000
14.750 1.1668 0.08840 0.08335 -0.0241 0.0077 1.0000
15.000 1.1526 0.09454 0.08971 -0.0268 0.0077 1.0000
15.250 1.1404 0.10047 0.09581 -0.0298 0.0077 1.0000
15.500 1.1267 0.10704 0.10255 -0.0333 0.0077 1.0000
15.750 1.1124 0.11410 0.10979 -0.0373 0.0076 1.0000
16.000 1.0974 0.12171 0.11758 -0.0419 0.0076 1.0000
16.250 1.0818 0.12993 0.12597 -0.0471 0.0077 1.0000
16.500 1.0393 0.14713 0.14356 -0.0584 0.0083 1.0000
16.750 1.0110 0.16192 0.15851 -0.0676 0.0086 1.0000
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Polar data table (+)
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