GOE 615 AIRFOIL (goe615-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 615 AIRFOIL (goe615-il) Reynolds number: 100,000 Max Cl/Cd: 52.96 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe615-il-100000-n5.txt Download as CSV file: xf-goe615-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 615 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.1398 0.09335 0.08844 -0.0696 0.9491 0.0843
-8.500 -0.1507 0.08880 0.08392 -0.0809 0.9356 0.0868
-8.000 -0.1444 0.07147 0.06645 -0.0932 0.9174 0.0571
-7.750 -0.1248 0.06774 0.06267 -0.0968 0.9115 0.0566
-7.500 -0.1169 0.06371 0.05859 -0.1001 0.9003 0.0561
-7.250 -0.1049 0.05785 0.05258 -0.1064 0.8922 0.0556
-7.000 -0.1047 0.05139 0.04586 -0.1109 0.8798 0.0558
-6.750 -0.1008 0.04608 0.04019 -0.1133 0.8690 0.0559
-6.500 -0.0884 0.04158 0.03526 -0.1150 0.8607 0.0558
-6.250 -0.0777 0.03829 0.03158 -0.1147 0.8506 0.0558
-6.000 -0.0586 0.03520 0.02800 -0.1152 0.8433 0.0560
-5.750 -0.0422 0.03289 0.02523 -0.1143 0.8342 0.0565
-5.500 -0.0198 0.03086 0.02277 -0.1141 0.8266 0.0576
-5.250 0.0036 0.02956 0.02135 -0.1138 0.8186 0.0587
-5.000 0.0279 0.02827 0.01986 -0.1134 0.8101 0.0595
-4.750 0.0535 0.02695 0.01827 -0.1130 0.8024 0.0602
-4.500 0.0777 0.02582 0.01691 -0.1124 0.7933 0.0610
-4.250 0.1053 0.02473 0.01558 -0.1122 0.7859 0.0619
-4.000 0.1295 0.02390 0.01456 -0.1114 0.7767 0.0634
-3.750 0.1589 0.02310 0.01346 -0.1114 0.7698 0.0655
-3.500 0.1823 0.02236 0.01268 -0.1104 0.7601 0.0669
-3.250 0.2120 0.02160 0.01186 -0.1106 0.7533 0.0684
-3.000 0.2349 0.02111 0.01133 -0.1095 0.7430 0.0698
-2.750 0.2642 0.02051 0.01064 -0.1094 0.7360 0.0719
-2.500 0.2872 0.02018 0.01024 -0.1083 0.7260 0.0747
-2.250 0.3158 0.01968 0.00971 -0.1083 0.7195 0.0780
-2.000 0.3384 0.01941 0.00946 -0.1072 0.7102 0.0812
-1.750 0.3662 0.01907 0.00902 -0.1069 0.7034 0.0852
-1.500 0.3898 0.01883 0.00878 -0.1059 0.6947 0.0902
-1.250 0.4167 0.01858 0.00847 -0.1055 0.6875 0.0990
-1.000 0.4416 0.01831 0.00823 -0.1048 0.6797 0.1114
-0.750 0.4667 0.01788 0.00804 -0.1042 0.6711 0.1522
-0.500 0.4916 0.01750 0.00798 -0.1036 0.6617 0.2620
-0.250 0.5185 0.01716 0.00785 -0.1032 0.6523 0.3546
0.000 0.5401 0.01676 0.00789 -0.1020 0.6419 0.4890
0.250 0.6211 0.01566 0.00778 -0.1116 0.6322 1.0000
0.500 0.6429 0.01584 0.00781 -0.1104 0.6238 1.0000
1.000 0.6900 0.01616 0.00782 -0.1084 0.6084 1.0000
1.250 0.7151 0.01630 0.00779 -0.1077 0.6014 1.0000
1.500 0.7372 0.01651 0.00790 -0.1065 0.5932 1.0000
1.750 0.7615 0.01668 0.00793 -0.1056 0.5858 1.0000
2.000 0.7847 0.01688 0.00804 -0.1047 0.5782 1.0000
2.250 0.8078 0.01708 0.00814 -0.1036 0.5701 1.0000
2.500 0.8318 0.01728 0.00823 -0.1028 0.5626 1.0000
2.750 0.8538 0.01751 0.00841 -0.1016 0.5541 1.0000
3.000 0.8782 0.01772 0.00850 -0.1008 0.5464 1.0000
3.250 0.8991 0.01797 0.00873 -0.0995 0.5373 1.0000
3.500 0.9234 0.01819 0.00884 -0.0987 0.5294 1.0000
3.750 0.9433 0.01846 0.00910 -0.0972 0.5197 1.0000
4.000 0.9663 0.01870 0.00926 -0.0962 0.5110 1.0000
4.250 0.9857 0.01898 0.00953 -0.0946 0.5009 1.0000
4.500 1.0066 0.01924 0.00975 -0.0933 0.4911 1.0000
4.750 1.0262 0.01951 0.01000 -0.0918 0.4806 1.0000
5.000 1.0443 0.01981 0.01032 -0.0900 0.4693 1.0000
5.250 1.0634 0.02009 0.01055 -0.0884 0.4586 1.0000
5.500 1.0803 0.02040 0.01085 -0.0865 0.4463 1.0000
5.750 1.0961 0.02073 0.01120 -0.0844 0.4333 1.0000
6.000 1.1112 0.02107 0.01151 -0.0822 0.4198 1.0000
6.250 1.1255 0.02143 0.01183 -0.0799 0.4063 1.0000
6.500 1.1379 0.02181 0.01215 -0.0772 0.3930 1.0000
7.000 1.1608 0.02277 0.01295 -0.0719 0.3672 1.0000
7.250 1.1728 0.02333 0.01345 -0.0695 0.3558 1.0000
7.500 1.1849 0.02395 0.01398 -0.0672 0.3456 1.0000
7.750 1.1977 0.02459 0.01460 -0.0651 0.3363 1.0000
8.000 1.2105 0.02527 0.01523 -0.0631 0.3278 1.0000
8.250 1.2222 0.02600 0.01593 -0.0611 0.3188 1.0000
8.500 1.2345 0.02675 0.01668 -0.0591 0.3107 1.0000
8.750 1.2458 0.02755 0.01748 -0.0572 0.3026 1.0000
9.000 1.2585 0.02835 0.01825 -0.0555 0.2960 1.0000
9.250 1.2712 0.02916 0.01914 -0.0539 0.2897 1.0000
9.500 1.2838 0.03000 0.01999 -0.0523 0.2840 1.0000
9.750 1.2966 0.03086 0.02088 -0.0508 0.2786 1.0000
10.000 1.3079 0.03180 0.02192 -0.0493 0.2727 1.0000
10.250 1.3193 0.03274 0.02288 -0.0478 0.2675 1.0000
10.500 1.3305 0.03374 0.02393 -0.0463 0.2622 1.0000
10.750 1.3408 0.03481 0.02513 -0.0449 0.2570 1.0000
11.000 1.3499 0.03594 0.02631 -0.0435 0.2518 1.0000
11.250 1.3590 0.03711 0.02750 -0.0421 0.2466 1.0000
11.500 1.3654 0.03849 0.02903 -0.0407 0.2407 1.0000
11.750 1.3728 0.03984 0.03046 -0.0394 0.2357 1.0000
12.000 1.3801 0.04122 0.03187 -0.0382 0.2310 1.0000
12.250 1.3856 0.04284 0.03366 -0.0370 0.2257 1.0000
12.500 1.3909 0.04447 0.03538 -0.0359 0.2209 1.0000
12.750 1.3961 0.04610 0.03702 -0.0349 0.2164 1.0000
13.000 1.3975 0.04823 0.03936 -0.0340 0.2102 1.0000
13.250 1.3985 0.05039 0.04161 -0.0332 0.2045 1.0000
13.500 1.3995 0.05262 0.04393 -0.0325 0.1991 1.0000
13.750 1.4002 0.05507 0.04653 -0.0319 0.1938 1.0000
14.000 1.3999 0.05762 0.04917 -0.0315 0.1886 1.0000
14.250 1.3991 0.06036 0.05203 -0.0312 0.1835 1.0000
14.500 1.3963 0.06345 0.05529 -0.0312 0.1778 1.0000
14.750 1.3926 0.06662 0.05851 -0.0312 0.1724 1.0000
15.000 1.3867 0.07038 0.06246 -0.0316 0.1660 1.0000
15.250 1.3796 0.07426 0.06642 -0.0320 0.1598 1.0000
15.500 1.3711 0.07853 0.07084 -0.0327 0.1534 1.0000
15.750 1.3603 0.08319 0.07558 -0.0337 0.1466 1.0000
16.000 1.3479 0.08827 0.08078 -0.0349 0.1396 1.0000
16.250 1.3338 0.09367 0.08623 -0.0364 0.1327 1.0000
16.500 1.3194 0.09933 0.09201 -0.0381 0.1257 1.0000
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Polar data table (+)
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