GOE 596 AIRFOIL (goe596-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 596 AIRFOIL (goe596-il) Reynolds number: 100,000 Max Cl/Cd: 56.55 at α=8° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe596-il-100000.txt Download as CSV file: xf-goe596-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 596 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3438 0.10358 0.09848 -0.0341 1.0000 0.0891
-8.750 -0.3637 0.10305 0.09810 -0.0371 1.0000 0.0900
-8.500 -0.3839 0.10196 0.09714 -0.0391 1.0000 0.0903
-8.250 -0.3402 0.09399 0.08907 -0.0337 1.0000 0.0934
-8.000 -0.3347 0.09133 0.08645 -0.0325 1.0000 0.0960
-7.750 -0.3376 0.08915 0.08435 -0.0317 1.0000 0.0986
-7.500 -0.3483 0.08740 0.08271 -0.0305 1.0000 0.1007
-7.250 -0.3648 0.08588 0.08131 -0.0299 1.0000 0.1027
-7.000 -0.3894 0.08487 0.08037 -0.0333 1.0000 0.1047
-6.750 -0.4075 0.08295 0.07840 -0.0368 1.0000 0.1057
-6.500 -0.3973 0.07869 0.07432 -0.0295 1.0000 0.1076
-6.250 -0.3949 0.07669 0.07239 -0.0255 1.0000 0.1104
-6.000 -0.3974 0.07467 0.07040 -0.0241 1.0000 0.1139
-5.750 -0.4054 0.07239 0.06783 -0.0329 1.0000 0.1213
-5.500 -0.4031 0.06888 0.06453 -0.0282 1.0000 0.1229
-5.250 -0.4012 0.06685 0.06258 -0.0250 1.0000 0.1253
-5.000 -0.3634 0.06221 0.05765 -0.0352 0.9924 0.1381
-4.750 -0.3305 0.05898 0.05448 -0.0374 0.9850 0.1456
-4.500 -0.2933 0.05513 0.05050 -0.0434 0.9753 0.1591
-4.250 -0.2524 0.05182 0.04688 -0.0510 0.9648 0.1848
-4.000 -0.2150 0.04854 0.04360 -0.0546 0.9573 0.2023
-3.750 -0.1832 0.04587 0.04090 -0.0570 0.9462 0.2212
-3.500 -0.1481 0.04329 0.03822 -0.0602 0.9363 0.2510
-3.250 -0.1120 0.04084 0.03578 -0.0626 0.9281 0.2850
-2.750 0.0095 0.02963 0.02193 -0.0762 0.9122 0.1188
-2.500 0.0495 0.02787 0.01947 -0.0773 0.9014 0.1055
-2.250 0.0969 0.02526 0.01665 -0.0807 0.8963 0.1026
-2.000 0.1323 0.02409 0.01522 -0.0816 0.8847 0.1035
-1.750 0.1794 0.02267 0.01354 -0.0844 0.8792 0.1043
-1.500 0.2113 0.02172 0.01247 -0.0846 0.8666 0.1048
-1.250 0.2445 0.02073 0.01144 -0.0849 0.8553 0.1063
-1.000 0.2834 0.01952 0.01033 -0.0863 0.8478 0.1113
-0.750 0.3108 0.01899 0.00979 -0.0856 0.8344 0.1193
-0.500 0.3374 0.01835 0.00923 -0.0847 0.8216 0.1284
-0.250 0.3660 0.01779 0.00870 -0.0842 0.8101 0.1463
0.000 0.4467 0.01471 0.00784 -0.0933 0.8025 1.0000
0.250 0.4704 0.01484 0.00776 -0.0921 0.7896 1.0000
0.500 0.4950 0.01497 0.00770 -0.0911 0.7777 1.0000
0.750 0.5220 0.01502 0.00757 -0.0904 0.7675 1.0000
1.000 0.5456 0.01519 0.00762 -0.0893 0.7555 1.0000
1.250 0.5690 0.01540 0.00772 -0.0882 0.7437 1.0000
1.500 0.5943 0.01555 0.00776 -0.0874 0.7333 1.0000
1.750 0.6203 0.01567 0.00776 -0.0866 0.7231 1.0000
2.000 0.6434 0.01592 0.00797 -0.0856 0.7114 1.0000
2.250 0.6682 0.01614 0.00811 -0.0847 0.7010 1.0000
2.500 0.6953 0.01627 0.00813 -0.0841 0.6914 1.0000
2.750 0.7183 0.01657 0.00842 -0.0831 0.6795 1.0000
3.000 0.7425 0.01686 0.00868 -0.0822 0.6686 1.0000
3.250 0.7694 0.01707 0.00880 -0.0817 0.6589 1.0000
3.500 0.7932 0.01739 0.00913 -0.0807 0.6471 1.0000
3.750 0.8166 0.01777 0.00951 -0.0798 0.6354 1.0000
4.000 0.8418 0.01811 0.00983 -0.0791 0.6246 1.0000
4.250 0.8683 0.01838 0.01002 -0.0784 0.6133 1.0000
4.500 0.8910 0.01872 0.01040 -0.0773 0.5995 1.0000
4.750 0.9141 0.01903 0.01072 -0.0761 0.5855 1.0000
5.000 0.9374 0.01936 0.01106 -0.0751 0.5721 1.0000
5.250 0.9615 0.01972 0.01145 -0.0742 0.5601 1.0000
5.500 0.9883 0.02002 0.01171 -0.0736 0.5495 1.0000
5.750 1.0100 0.02039 0.01218 -0.0724 0.5361 1.0000
6.000 1.0326 0.02060 0.01242 -0.0711 0.5214 1.0000
6.250 1.0555 0.02063 0.01247 -0.0698 0.5053 1.0000
6.500 1.0781 0.02059 0.01241 -0.0684 0.4884 1.0000
6.750 1.1001 0.02064 0.01249 -0.0669 0.4722 1.0000
7.000 1.1217 0.02068 0.01258 -0.0655 0.4558 1.0000
7.250 1.1407 0.02079 0.01279 -0.0637 0.4380 1.0000
7.500 1.1592 0.02088 0.01299 -0.0618 0.4188 1.0000
7.750 1.1774 0.02095 0.01311 -0.0599 0.3983 1.0000
8.000 1.1927 0.02109 0.01333 -0.0575 0.3728 1.0000
8.250 1.2059 0.02139 0.01362 -0.0548 0.3429 1.0000
8.500 1.2144 0.02194 0.01409 -0.0516 0.3019 1.0000
8.750 1.2132 0.02312 0.01495 -0.0471 0.2252 1.0000
9.000 1.2009 0.02537 0.01652 -0.0418 0.1553 1.0000
9.250 1.1944 0.02728 0.01815 -0.0372 0.1265 1.0000
9.500 1.1932 0.02895 0.01971 -0.0335 0.1124 1.0000
9.750 1.1924 0.03067 0.02136 -0.0303 0.1046 1.0000
10.000 1.1957 0.03227 0.02300 -0.0277 0.0978 1.0000
10.250 1.1972 0.03414 0.02478 -0.0252 0.0927 1.0000
10.500 1.2054 0.03572 0.02644 -0.0233 0.0879 1.0000
10.750 1.2155 0.03741 0.02808 -0.0216 0.0838 1.0000
11.000 1.2343 0.03914 0.02976 -0.0204 0.0796 1.0000
11.250 1.2514 0.04078 0.03152 -0.0193 0.0755 1.0000
11.500 1.2827 0.04271 0.03336 -0.0193 0.0718 1.0000
11.750 1.3143 0.04539 0.03616 -0.0197 0.0684 1.0000
12.000 1.3278 0.04760 0.03864 -0.0184 0.0662 1.0000
12.250 1.3446 0.05026 0.04157 -0.0175 0.0645 1.0000
12.500 1.3581 0.05302 0.04453 -0.0165 0.0629 1.0000
12.750 1.3694 0.05598 0.04763 -0.0157 0.0610 1.0000
13.000 1.3892 0.06180 0.05360 -0.0165 0.0591 1.0000
13.250 1.3776 0.06463 0.05675 -0.0137 0.0589 1.0000
13.500 1.3655 0.06798 0.06040 -0.0115 0.0589 1.0000
13.750 1.3528 0.07180 0.06449 -0.0099 0.0589 1.0000
14.000 1.3387 0.07604 0.06899 -0.0088 0.0590 1.0000
14.250 1.3222 0.08055 0.07378 -0.0083 0.0592 1.0000
14.500 1.3004 0.08436 0.07784 -0.0082 0.0594 1.0000
14.750 1.2746 0.08884 0.08259 -0.0091 0.0598 1.0000
15.000 1.2374 0.09494 0.08902 -0.0119 0.0606 1.0000
15.250 1.1550 0.10870 0.10326 -0.0216 0.0631 1.0000
15.500 1.0842 0.12674 0.12153 -0.0338 0.0671 1.0000
15.750 1.0545 0.13892 0.13376 -0.0408 0.0690 1.0000
16.000 1.0394 0.14869 0.14356 -0.0457 0.0705 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 596 AIRFOIL (goe596-il)