GOE 626 AIRFOIL (goe626-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 626 AIRFOIL (goe626-il) Reynolds number: 50,000 Max Cl/Cd: 30.55 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe626-il-50000-n5.txt Download as CSV file: xf-goe626-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 626 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
------ -------- --------- --------- -------- -------- --------
-11.000 -0.3385 0.11011 0.10227 -0.0483 1.0000 0.0913
-10.750 -0.3562 0.10452 0.09675 -0.0494 1.0000 0.0924
-10.500 -0.3749 0.09913 0.09145 -0.0502 1.0000 0.0938
-10.250 -0.3720 0.09810 0.09048 -0.0485 1.0000 0.0955
-10.000 -0.3754 0.09613 0.08859 -0.0473 1.0000 0.0978
-9.750 -0.3948 0.09171 0.08426 -0.0473 1.0000 0.1004
-9.500 -0.4768 0.07856 0.07124 -0.0509 1.0000 0.1038
-9.250 -0.4507 0.08168 0.07445 -0.0470 1.0000 0.1063
-9.000 -0.4567 0.08051 0.07338 -0.0449 1.0000 0.1093
-8.750 -0.5108 0.07397 0.06697 -0.0443 1.0000 0.1122
-8.500 -0.5841 0.06497 0.05803 -0.0455 1.0000 0.1136
-8.250 -0.5649 0.06537 0.05855 -0.0449 0.9973 0.1190
-8.000 -0.5538 0.06003 0.05316 -0.0528 0.9875 0.1292
-7.750 -0.5315 0.05632 0.04936 -0.0596 0.9785 0.1407
-7.250 -0.4524 0.06044 0.05342 -0.0597 0.9633 0.1581
-6.750 -0.3909 0.05967 0.05248 -0.0657 0.9443 0.1843
-6.500 -0.3571 0.05884 0.05152 -0.0696 0.9360 0.1980
-6.250 -0.3342 0.05610 0.04861 -0.0745 0.9245 0.2109
-6.000 -0.3096 0.05293 0.04522 -0.0799 0.9138 0.2243
-5.750 -0.2740 0.05408 0.04634 -0.0789 0.9057 0.2284
-5.500 -0.2507 0.05047 0.04246 -0.0843 0.8939 0.2408
-5.000 -0.1902 0.04805 0.03973 -0.0886 0.8739 0.2579
-4.750 -0.1490 0.04558 0.03692 -0.0946 0.8671 0.2717
-4.500 -0.1283 0.04591 0.03730 -0.0922 0.8540 0.2756
-4.250 -0.0852 0.04395 0.03502 -0.0975 0.8479 0.2894
-4.000 -0.0663 0.04407 0.03518 -0.0952 0.8343 0.2936
-3.750 -0.0209 0.04203 0.03278 -0.1008 0.8287 0.3070
-3.500 -0.0032 0.04194 0.03274 -0.0985 0.8150 0.3107
-3.250 0.0324 0.04110 0.03179 -0.1000 0.8073 0.3179
-3.000 0.0628 0.03955 0.02992 -0.1027 0.7958 0.3264
-2.750 0.0932 0.03906 0.02943 -0.1027 0.7868 0.3307
-2.500 0.1234 0.03828 0.02854 -0.1035 0.7765 0.3365
-2.250 0.1577 0.03702 0.02694 -0.1063 0.7665 0.3438
-2.000 0.1894 0.03634 0.02623 -0.1068 0.7572 0.3478
-1.750 0.2156 0.03588 0.02572 -0.1066 0.7463 0.3518
-1.500 0.2520 0.03501 0.02469 -0.1082 0.7379 0.3568
-1.250 0.2804 0.03438 0.02383 -0.1091 0.7263 0.3621
-1.000 0.3182 0.03357 0.02290 -0.1107 0.7187 0.3665
-0.750 0.3420 0.03331 0.02260 -0.1101 0.7069 0.3695
-0.500 0.3814 0.03256 0.02172 -0.1118 0.6997 0.3737
-0.250 0.4059 0.03234 0.02139 -0.1116 0.6876 0.3781
0.000 0.4482 0.03158 0.02035 -0.1141 0.6807 0.3834
0.250 0.4705 0.03152 0.02030 -0.1132 0.6688 0.3860
0.500 0.5077 0.03102 0.01969 -0.1144 0.6612 0.3894
0.750 0.5317 0.03103 0.01967 -0.1139 0.6500 0.3932
1.000 0.5705 0.03061 0.01907 -0.1155 0.6428 0.3983
1.250 0.5933 0.03077 0.01913 -0.1151 0.6315 0.4024
1.500 0.6304 0.03044 0.01872 -0.1162 0.6244 0.4059
1.750 0.6498 0.03075 0.01906 -0.1151 0.6136 0.4087
2.000 0.6843 0.03062 0.01884 -0.1159 0.6062 0.4132
2.250 0.7069 0.03097 0.01915 -0.1153 0.5965 0.4178
2.500 0.7376 0.03107 0.01914 -0.1158 0.5884 0.4228
2.750 0.7649 0.03127 0.01935 -0.1156 0.5807 0.4270
3.000 0.7867 0.03169 0.01980 -0.1148 0.5718 0.4313
3.250 0.8221 0.03168 0.01968 -0.1158 0.5655 0.4366
3.500 0.8374 0.03240 0.02043 -0.1142 0.5562 0.4409
3.750 0.8649 0.03263 0.02071 -0.1141 0.5492 0.4462
4.000 0.8928 0.03293 0.02101 -0.1141 0.5429 0.4529
4.250 0.9075 0.03372 0.02188 -0.1124 0.5344 0.4592
4.500 0.9382 0.03385 0.02205 -0.1127 0.5283 0.4663
4.750 0.9552 0.03457 0.02285 -0.1113 0.5209 0.4731
5.000 0.9743 0.03518 0.02353 -0.1101 0.5133 0.4805
5.250 1.0104 0.03501 0.02338 -0.1109 0.5076 0.4910
5.500 1.0134 0.03622 0.02477 -0.1078 0.4981 0.4992
5.750 1.0451 0.03605 0.02464 -0.1079 0.4908 0.5143
6.000 1.0539 0.03689 0.02568 -0.1054 0.4816 0.5270
6.250 1.0813 0.03674 0.02565 -0.1049 0.4732 0.5492
6.500 1.0920 0.03735 0.02647 -0.1025 0.4639 0.5742
7.000 1.1235 0.03724 0.02712 -0.0981 0.4484 1.0000
7.250 1.1308 0.03835 0.02822 -0.0956 0.4403 1.0000
7.500 1.1690 0.03826 0.02796 -0.0965 0.4345 1.0000
7.750 1.1553 0.04033 0.03016 -0.0918 0.4261 1.0000
8.000 1.1772 0.04083 0.03062 -0.0909 0.4192 1.0000
8.250 1.1909 0.04173 0.03152 -0.0892 0.4122 1.0000
8.500 1.1855 0.04364 0.03352 -0.0859 0.4038 1.0000
8.750 1.2284 0.04294 0.03272 -0.0867 0.3977 1.0000
9.000 1.1936 0.04669 0.03667 -0.0815 0.3881 1.0000
9.250 1.2238 0.04655 0.03649 -0.0810 0.3814 1.0000
9.500 1.2009 0.05009 0.04017 -0.0777 0.3722 1.0000
9.750 1.2129 0.05125 0.04136 -0.0763 0.3647 1.0000
10.000 1.2144 0.05333 0.04350 -0.0747 0.3570 1.0000
10.250 1.1973 0.05713 0.04742 -0.0729 0.3472 1.0000
10.500 1.2528 0.05441 0.04460 -0.0726 0.3425 1.0000
10.750 1.1820 0.06380 0.05425 -0.0708 0.3288 1.0000
11.000 1.2159 0.06263 0.05304 -0.0696 0.3233 1.0000
11.250 1.1693 0.07092 0.06149 -0.0698 0.3104 1.0000
11.750 1.1551 0.07897 0.06965 -0.0699 0.2924 1.0000
12.250 1.1354 0.08862 0.07942 -0.0712 0.2752 1.0000
12.750 1.1264 0.09672 0.08762 -0.0724 0.2586 1.0000
13.000 1.0782 0.10915 0.10015 -0.0770 0.2458 1.0000
13.250 1.1283 0.10290 0.09389 -0.0732 0.2425 1.0000
13.750 1.1367 0.10784 0.09890 -0.0737 0.2268 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 626 AIRFOIL (goe626-il)