GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 50,000 Max Cl/Cd: 36.47 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe566-il-50000-n5.txt Download as CSV file: xf-goe566-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 566 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4354 0.09270 0.08568 -0.0317 1.0000 0.0821
-8.000 -0.4615 0.08276 0.07588 -0.0402 1.0000 0.0705
-7.750 -0.4612 0.07899 0.07212 -0.0405 1.0000 0.0703
-7.500 -0.4613 0.07529 0.06842 -0.0407 1.0000 0.0701
-7.250 -0.4608 0.07162 0.06475 -0.0406 1.0000 0.0697
-7.000 -0.4595 0.06806 0.06117 -0.0401 1.0000 0.0690
-6.750 -0.4584 0.06446 0.05751 -0.0397 1.0000 0.0683
-6.500 -0.4569 0.06082 0.05377 -0.0392 1.0000 0.0676
-6.250 -0.4543 0.05718 0.04995 -0.0386 1.0000 0.0678
-6.000 -0.4503 0.05352 0.04600 -0.0379 1.0000 0.0688
-5.750 -0.4442 0.04989 0.04196 -0.0370 1.0000 0.0699
-5.500 -0.4350 0.04658 0.03826 -0.0359 1.0000 0.0702
-5.250 -0.4233 0.04354 0.03482 -0.0346 1.0000 0.0702
-5.000 -0.4097 0.04072 0.03158 -0.0333 1.0000 0.0705
-4.750 -0.3951 0.03836 0.02903 -0.0322 1.0000 0.0721
-4.500 -0.3794 0.03687 0.02747 -0.0310 1.0000 0.0751
-4.250 -0.3619 0.03500 0.02528 -0.0298 1.0000 0.0773
-4.000 -0.3428 0.03302 0.02290 -0.0287 1.0000 0.0782
-3.750 -0.3226 0.03126 0.02073 -0.0276 1.0000 0.0792
-3.500 -0.3016 0.02974 0.01884 -0.0266 1.0000 0.0805
-3.250 -0.2801 0.02846 0.01722 -0.0256 1.0000 0.0823
-3.000 -0.2588 0.02745 0.01599 -0.0246 1.0000 0.0859
-2.750 -0.2288 0.02660 0.01504 -0.0255 0.9961 0.0906
-2.500 -0.1938 0.02568 0.01391 -0.0268 0.9904 0.0943
-2.250 -0.1573 0.02495 0.01291 -0.0284 0.9852 0.0985
-2.000 -0.1239 0.02427 0.01222 -0.0297 0.9788 0.1042
-1.750 -0.0890 0.02380 0.01159 -0.0312 0.9726 0.1139
-1.500 -0.0558 0.02331 0.01119 -0.0327 0.9658 0.1338
-1.250 -0.0226 0.02263 0.01072 -0.0342 0.9591 0.1773
-1.000 0.0344 0.01955 0.01055 -0.0391 0.9600 1.0000
-0.750 0.0691 0.01981 0.01039 -0.0408 0.9515 1.0000
-0.500 0.1021 0.02007 0.01035 -0.0422 0.9426 1.0000
-0.250 0.1407 0.02035 0.01037 -0.0446 0.9347 1.0000
0.000 0.1729 0.02057 0.01040 -0.0457 0.9238 1.0000
0.250 0.2088 0.02078 0.01044 -0.0475 0.9132 1.0000
0.500 0.2517 0.02094 0.01046 -0.0504 0.9041 1.0000
0.750 0.2851 0.02109 0.01051 -0.0516 0.8918 1.0000
1.000 0.3200 0.02120 0.01055 -0.0528 0.8793 1.0000
1.250 0.3584 0.02113 0.01042 -0.0543 0.8642 1.0000
1.500 0.3979 0.02090 0.01015 -0.0555 0.8472 1.0000
1.750 0.4314 0.02070 0.00993 -0.0557 0.8290 1.0000
2.000 0.4597 0.02062 0.00984 -0.0551 0.8109 1.0000
2.250 0.4882 0.02057 0.00981 -0.0545 0.7945 1.0000
2.500 0.5160 0.02054 0.00981 -0.0539 0.7780 1.0000
2.750 0.5437 0.02049 0.00979 -0.0532 0.7608 1.0000
3.000 0.5720 0.02040 0.00973 -0.0524 0.7426 1.0000
3.250 0.5965 0.02041 0.00981 -0.0512 0.7214 1.0000
3.500 0.6232 0.02035 0.00979 -0.0501 0.6992 1.0000
3.750 0.6481 0.02036 0.00984 -0.0488 0.6737 1.0000
4.000 0.6737 0.02038 0.00989 -0.0476 0.6459 1.0000
4.250 0.6987 0.02045 0.00996 -0.0463 0.6148 1.0000
4.500 0.7229 0.02060 0.01008 -0.0448 0.5801 1.0000
4.750 0.7459 0.02086 0.01031 -0.0433 0.5427 1.0000
5.000 0.7687 0.02122 0.01057 -0.0419 0.5063 1.0000
5.250 0.7912 0.02170 0.01088 -0.0405 0.4732 1.0000
5.500 0.8130 0.02229 0.01134 -0.0392 0.4434 1.0000
5.750 0.8344 0.02297 0.01196 -0.0380 0.4165 1.0000
6.000 0.8555 0.02369 0.01262 -0.0368 0.3916 1.0000
6.250 0.8761 0.02442 0.01327 -0.0355 0.3680 1.0000
6.500 0.8962 0.02515 0.01405 -0.0343 0.3454 1.0000
6.750 0.9170 0.02590 0.01486 -0.0332 0.3263 1.0000
7.000 0.9387 0.02667 0.01572 -0.0322 0.3102 1.0000
7.250 0.9605 0.02750 0.01665 -0.0313 0.2946 1.0000
7.500 0.9804 0.02836 0.01765 -0.0301 0.2769 1.0000
7.750 0.9967 0.02922 0.01863 -0.0284 0.2549 1.0000
8.000 1.0098 0.03010 0.01958 -0.0265 0.2308 1.0000
8.250 1.0205 0.03099 0.02051 -0.0243 0.2056 1.0000
8.500 1.0313 0.03196 0.02163 -0.0222 0.1792 1.0000
8.750 1.0421 0.03305 0.02280 -0.0202 0.1542 1.0000
9.000 1.0507 0.03438 0.02405 -0.0180 0.1338 1.0000
9.250 1.0584 0.03603 0.02565 -0.0158 0.1141 1.0000
9.500 1.0634 0.03791 0.02748 -0.0135 0.0995 1.0000
9.750 1.0701 0.03999 0.02959 -0.0113 0.0889 1.0000
10.000 1.0759 0.04216 0.03173 -0.0093 0.0806 1.0000
10.250 1.0840 0.04435 0.03414 -0.0075 0.0730 1.0000
10.500 1.0910 0.04666 0.03648 -0.0059 0.0678 1.0000
10.750 1.0962 0.04903 0.03921 -0.0042 0.0625 1.0000
11.000 1.1000 0.05135 0.04162 -0.0027 0.0588 1.0000
11.250 1.1050 0.05409 0.04453 -0.0014 0.0562 1.0000
11.500 1.1046 0.05725 0.04806 -0.0001 0.0538 1.0000
11.750 1.1011 0.06046 0.05153 0.0008 0.0516 1.0000
12.000 1.0970 0.06365 0.05489 0.0013 0.0497 1.0000
12.250 1.0935 0.06682 0.05821 0.0015 0.0482 1.0000
12.500 1.0872 0.07075 0.06226 0.0014 0.0470 1.0000
12.750 1.0720 0.07596 0.06777 0.0002 0.0466 1.0000
13.000 1.0552 0.08174 0.07381 -0.0019 0.0465 1.0000
13.250 1.0359 0.08834 0.08064 -0.0050 0.0465 1.0000
13.500 1.0143 0.09594 0.08838 -0.0092 0.0467 1.0000
13.750 0.9920 0.10449 0.09709 -0.0142 0.0469 1.0000
14.000 0.9704 0.11367 0.10640 -0.0197 0.0473 1.0000
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Polar data table (+)
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