GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 200,000 Max Cl/Cd: 69.97 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe566-il-200000.txt Download as CSV file: xf-goe566-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3585 0.09529 0.09182 -0.0309 1.0000 0.0519
-9.500 -0.3675 0.09144 0.08801 -0.0342 1.0000 0.0522
-9.250 -0.3740 0.08739 0.08401 -0.0369 1.0000 0.0524
-9.000 -0.3670 0.08133 0.07796 -0.0347 1.0000 0.0536
-8.750 -0.3546 0.07848 0.07512 -0.0327 1.0000 0.0550
-8.500 -0.3505 0.07539 0.07204 -0.0324 1.0000 0.0565
-8.250 -0.3518 0.07194 0.06863 -0.0328 1.0000 0.0581
-8.000 -0.4620 0.07634 0.07294 -0.0384 1.0000 0.0540
-7.750 -0.4567 0.07415 0.07077 -0.0364 1.0000 0.0552
-7.500 -0.4563 0.07149 0.06814 -0.0359 1.0000 0.0566
-7.250 -0.4596 0.06857 0.06523 -0.0357 1.0000 0.0578
-7.000 -0.4654 0.06568 0.06234 -0.0350 1.0000 0.0594
-6.750 -0.4725 0.06270 0.05931 -0.0342 1.0000 0.0615
-6.500 -0.4858 0.06027 0.05630 -0.0346 1.0000 0.0649
-6.250 -0.4875 0.05456 0.05079 -0.0329 1.0000 0.0665
-6.000 -0.4814 0.05295 0.04926 -0.0305 1.0000 0.0680
-5.750 -0.4755 0.05117 0.04747 -0.0286 1.0000 0.0706
-5.500 -0.4545 0.04610 0.04179 -0.0321 0.9967 0.0794
-5.250 -0.4262 0.04323 0.03902 -0.0342 0.9933 0.0822
-5.000 -0.3947 0.03989 0.03516 -0.0373 0.9876 0.0936
-4.750 -0.3639 0.03757 0.03296 -0.0394 0.9843 0.0984
-4.500 -0.3344 0.03482 0.02989 -0.0413 0.9786 0.1100
-4.250 -0.2988 0.02571 0.01909 -0.0407 0.9740 0.0657
-4.000 -0.2620 0.02357 0.01667 -0.0425 0.9709 0.0640
-3.750 -0.2322 0.02140 0.01420 -0.0430 0.9658 0.0630
-3.500 -0.1977 0.02006 0.01258 -0.0442 0.9610 0.0637
-3.250 -0.1586 0.01919 0.01147 -0.0462 0.9576 0.0649
-3.000 -0.1233 0.01769 0.00983 -0.0476 0.9538 0.0655
-2.750 -0.0926 0.01656 0.00867 -0.0482 0.9479 0.0670
-2.500 -0.0547 0.01576 0.00790 -0.0502 0.9441 0.0696
-2.250 -0.0132 0.01516 0.00732 -0.0529 0.9415 0.0743
-2.000 0.0157 0.01472 0.00686 -0.0530 0.9341 0.0772
-1.750 0.0537 0.01401 0.00623 -0.0550 0.9297 0.0838
-1.500 0.0979 0.01332 0.00561 -0.0579 0.9260 0.0997
-1.250 0.1220 0.01197 0.00515 -0.0572 0.9146 0.2971
-1.000 0.1762 0.00983 0.00509 -0.0608 0.9120 0.9191
-0.750 0.2905 0.00954 0.00465 -0.0775 0.9136 1.0000
-0.500 0.3150 0.00940 0.00440 -0.0766 0.9002 1.0000
-0.250 0.3379 0.00930 0.00420 -0.0753 0.8872 1.0000
0.000 0.3608 0.00924 0.00406 -0.0740 0.8753 1.0000
0.250 0.3841 0.00919 0.00392 -0.0728 0.8639 1.0000
0.500 0.4065 0.00916 0.00382 -0.0713 0.8517 1.0000
0.750 0.4284 0.00915 0.00375 -0.0699 0.8387 1.0000
1.000 0.4509 0.00916 0.00370 -0.0685 0.8260 1.0000
1.250 0.4739 0.00916 0.00365 -0.0672 0.8132 1.0000
1.500 0.4971 0.00917 0.00361 -0.0659 0.7996 1.0000
1.750 0.5201 0.00916 0.00355 -0.0646 0.7837 1.0000
2.000 0.5432 0.00916 0.00349 -0.0632 0.7660 1.0000
2.250 0.5666 0.00918 0.00344 -0.0618 0.7474 1.0000
2.500 0.5895 0.00923 0.00344 -0.0605 0.7263 1.0000
2.750 0.6127 0.00930 0.00344 -0.0592 0.7046 1.0000
3.000 0.6355 0.00940 0.00346 -0.0578 0.6794 1.0000
3.250 0.6578 0.00954 0.00351 -0.0564 0.6500 1.0000
3.500 0.6795 0.00973 0.00357 -0.0548 0.6143 1.0000
3.750 0.6997 0.01000 0.00366 -0.0530 0.5655 1.0000
4.000 0.7181 0.01041 0.00378 -0.0509 0.4998 1.0000
4.250 0.7352 0.01099 0.00402 -0.0488 0.4392 1.0000
4.500 0.7539 0.01158 0.00436 -0.0471 0.3975 1.0000
4.750 0.7743 0.01210 0.00471 -0.0457 0.3681 1.0000
5.000 0.7953 0.01259 0.00507 -0.0445 0.3457 1.0000
5.250 0.8173 0.01300 0.00542 -0.0434 0.3250 1.0000
5.500 0.8393 0.01340 0.00577 -0.0424 0.3074 1.0000
5.750 0.8616 0.01379 0.00614 -0.0414 0.2926 1.0000
6.000 0.8840 0.01419 0.00652 -0.0405 0.2796 1.0000
6.250 0.9061 0.01457 0.00692 -0.0395 0.2649 1.0000
6.500 0.9280 0.01495 0.00730 -0.0385 0.2490 1.0000
6.750 0.9498 0.01533 0.00770 -0.0375 0.2330 1.0000
7.000 0.9709 0.01573 0.00809 -0.0364 0.2142 1.0000
7.250 0.9927 0.01606 0.00848 -0.0354 0.1901 1.0000
7.500 1.0130 0.01653 0.00886 -0.0343 0.1504 1.0000
7.750 1.0266 0.01773 0.00966 -0.0323 0.1008 1.0000
8.000 1.0372 0.01928 0.01103 -0.0297 0.0751 1.0000
8.250 1.0525 0.02033 0.01210 -0.0278 0.0620 1.0000
8.500 1.0685 0.02126 0.01301 -0.0260 0.0543 1.0000
8.750 1.0838 0.02225 0.01407 -0.0241 0.0496 1.0000
9.000 1.0982 0.02326 0.01509 -0.0222 0.0457 1.0000
9.250 1.1096 0.02470 0.01656 -0.0199 0.0425 1.0000
9.500 1.1241 0.02591 0.01787 -0.0179 0.0401 1.0000
9.750 1.1381 0.02723 0.01927 -0.0161 0.0382 1.0000
10.000 1.1519 0.02884 0.02090 -0.0143 0.0364 1.0000
10.250 1.1686 0.03131 0.02348 -0.0133 0.0344 1.0000
10.500 1.1826 0.03272 0.02509 -0.0115 0.0331 1.0000
10.750 1.1969 0.03457 0.02714 -0.0099 0.0320 1.0000
11.000 1.2094 0.03660 0.02938 -0.0081 0.0311 1.0000
11.250 1.2194 0.03874 0.03172 -0.0062 0.0304 1.0000
11.500 1.2260 0.04096 0.03415 -0.0040 0.0298 1.0000
11.750 1.2302 0.04315 0.03650 -0.0018 0.0290 1.0000
12.000 1.2335 0.04584 0.03934 0.0001 0.0283 1.0000
12.250 1.2313 0.04932 0.04303 0.0021 0.0278 1.0000
12.500 1.2201 0.05346 0.04745 0.0045 0.0275 1.0000
12.750 1.2040 0.05758 0.05185 0.0065 0.0273 1.0000
13.000 1.1888 0.06124 0.05574 0.0079 0.0273 1.0000
13.250 1.1712 0.06530 0.06004 0.0086 0.0273 1.0000
13.500 1.1575 0.06893 0.06386 0.0086 0.0274 1.0000
13.750 1.1392 0.07355 0.06869 0.0076 0.0274 1.0000
14.000 1.1214 0.07841 0.07375 0.0058 0.0275 1.0000
14.250 1.1035 0.08366 0.07919 0.0030 0.0277 1.0000
14.500 1.0818 0.09021 0.08594 -0.0011 0.0279 1.0000
14.750 1.0448 0.10087 0.09690 -0.0091 0.0286 1.0000
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