GOE 646 AIRFOIL (goe646-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 646 AIRFOIL (goe646-il) Reynolds number: 200,000 Max Cl/Cd: 68.84 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe646-il-200000.txt Download as CSV file: xf-goe646-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 646 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
------ -------- --------- --------- -------- -------- --------
-15.000 -0.7277 0.07265 0.06779 -0.0922 1.0000 0.0692
-14.750 -0.7842 0.06445 0.05938 -0.0960 1.0000 0.0690
-14.500 -0.8247 0.05933 0.05412 -0.0969 1.0000 0.0690
-14.250 -0.8639 0.05528 0.04993 -0.0965 1.0000 0.0690
-14.000 -0.8996 0.05206 0.04658 -0.0952 1.0000 0.0691
-13.750 -0.9325 0.04941 0.04381 -0.0933 1.0000 0.0691
-13.500 -0.9644 0.04722 0.04149 -0.0906 1.0000 0.0692
-13.250 -0.9899 0.04508 0.03915 -0.0892 0.9974 0.0694
-13.000 -0.9605 0.04324 0.03735 -0.0917 0.9938 0.0707
-12.750 -0.9373 0.04180 0.03589 -0.0938 0.9886 0.0720
-12.500 -0.9171 0.04000 0.03395 -0.0962 0.9828 0.0736
-12.250 -0.9010 0.03818 0.03189 -0.0979 0.9770 0.0753
-12.000 -0.8862 0.03632 0.02969 -0.0990 0.9702 0.0768
-11.750 -0.8510 0.03501 0.02852 -0.1013 0.9678 0.0789
-11.500 -0.8325 0.03394 0.02742 -0.1011 0.9616 0.0806
-11.250 -0.8085 0.03261 0.02590 -0.1023 0.9563 0.0831
-11.000 -0.7768 0.03113 0.02430 -0.1045 0.9533 0.0859
-10.750 -0.7591 0.03037 0.02361 -0.1036 0.9465 0.0882
-10.500 -0.7320 0.02937 0.02248 -0.1046 0.9416 0.0916
-10.250 -0.6983 0.02814 0.02121 -0.1067 0.9387 0.0954
-10.000 -0.6778 0.02745 0.02054 -0.1061 0.9323 0.0987
-9.750 -0.6518 0.02661 0.01949 -0.1065 0.9266 0.1029
-9.500 -0.6166 0.02565 0.01868 -0.1085 0.9236 0.1076
-9.250 -0.5788 0.02481 0.01768 -0.1109 0.9214 0.1135
-9.000 -0.5661 0.02428 0.01724 -0.1086 0.9124 0.1172
-8.750 -0.5315 0.02362 0.01647 -0.1102 0.9084 0.1232
-8.500 -0.4930 0.02276 0.01569 -0.1125 0.9059 0.1294
-8.250 -0.4511 0.02202 0.01484 -0.1153 0.9042 0.1368
-7.750 -0.4009 0.02102 0.01381 -0.1146 0.8910 0.1495
-7.500 -0.3600 0.02031 0.01318 -0.1172 0.8888 0.1585
-7.250 -0.3157 0.01960 0.01250 -0.1205 0.8870 0.1692
-7.000 -0.2709 0.01894 0.01185 -0.1238 0.8852 0.1819
-6.750 -0.2590 0.01879 0.01165 -0.1206 0.8737 0.1907
-6.500 -0.2184 0.01824 0.01115 -0.1230 0.8701 0.2043
-6.250 -0.1755 0.01777 0.01070 -0.1257 0.8673 0.2184
-6.000 -0.1561 0.01760 0.01052 -0.1239 0.8576 0.2286
-5.750 -0.1207 0.01735 0.01018 -0.1251 0.8518 0.2413
-5.500 -0.0792 0.01701 0.00989 -0.1275 0.8479 0.2531
-5.250 -0.0591 0.01689 0.00972 -0.1258 0.8375 0.2623
-5.000 -0.0239 0.01667 0.00950 -0.1268 0.8310 0.2725
-4.750 0.0095 0.01643 0.00920 -0.1276 0.8241 0.2824
-4.500 0.0336 0.01634 0.00911 -0.1266 0.8138 0.2911
-4.250 0.0714 0.01607 0.00874 -0.1282 0.8081 0.3009
-4.000 0.0911 0.01604 0.00874 -0.1264 0.7967 0.3084
-3.750 0.1254 0.01584 0.00840 -0.1273 0.7897 0.3181
-3.500 0.1483 0.01577 0.00838 -0.1261 0.7795 0.3254
-3.250 0.1794 0.01570 0.00818 -0.1264 0.7715 0.3356
-3.000 0.2047 0.01559 0.00813 -0.1257 0.7621 0.3440
-2.750 0.2331 0.01555 0.00801 -0.1255 0.7534 0.3538
-2.500 0.2613 0.01540 0.00783 -0.1254 0.7456 0.3616
-2.250 0.2864 0.01535 0.00777 -0.1247 0.7367 0.3691
-2.000 0.3193 0.01531 0.00755 -0.1255 0.7306 0.3777
-1.750 0.3406 0.01518 0.00752 -0.1241 0.7215 0.3844
-1.500 0.3697 0.01511 0.00739 -0.1241 0.7143 0.3927
-1.250 0.3961 0.01506 0.00727 -0.1237 0.7064 0.4005
-1.000 0.4214 0.01495 0.00720 -0.1230 0.6981 0.4079
-0.750 0.4524 0.01492 0.00704 -0.1233 0.6910 0.4171
-0.500 0.4737 0.01484 0.00701 -0.1219 0.6820 0.4246
-0.250 0.5033 0.01477 0.00689 -0.1220 0.6746 0.4338
0.000 0.5273 0.01480 0.00689 -0.1211 0.6671 0.4429
0.250 0.5526 0.01476 0.00691 -0.1205 0.6602 0.4523
0.500 0.5836 0.01479 0.00684 -0.1210 0.6545 0.4633
0.750 0.6059 0.01480 0.00697 -0.1198 0.6481 0.4733
1.000 0.6311 0.01484 0.00702 -0.1192 0.6419 0.4854
1.250 0.6602 0.01485 0.00705 -0.1193 0.6362 0.4985
1.500 0.6843 0.01491 0.00719 -0.1184 0.6302 0.5121
1.750 0.7074 0.01496 0.00731 -0.1174 0.6236 0.5286
2.000 0.7351 0.01501 0.00739 -0.1172 0.6180 0.5484
2.250 0.7615 0.01512 0.00757 -0.1168 0.6125 0.5704
2.500 0.7812 0.01518 0.00778 -0.1151 0.6058 0.5948
2.750 0.8065 0.01522 0.00790 -0.1143 0.5996 0.6250
3.000 0.8337 0.01534 0.00806 -0.1140 0.5938 0.6584
3.250 0.8507 0.01539 0.00827 -0.1117 0.5871 0.6903
3.500 0.8738 0.01537 0.00835 -0.1105 0.5806 0.7233
3.750 0.9008 0.01539 0.00844 -0.1100 0.5744 0.7613
4.000 0.9164 0.01533 0.00864 -0.1073 0.5669 0.8075
4.250 0.9627 0.01523 0.00869 -0.1107 0.5584 0.9039
4.500 1.0044 0.01537 0.00884 -0.1138 0.5479 1.0000
4.750 1.0273 0.01547 0.00883 -0.1127 0.5393 1.0000
5.000 1.0457 0.01565 0.00898 -0.1109 0.5307 1.0000
5.250 1.0647 0.01578 0.00907 -0.1091 0.5221 1.0000
5.500 1.0857 0.01596 0.00919 -0.1078 0.5136 1.0000
5.750 1.1002 0.01611 0.00933 -0.1051 0.5040 1.0000
6.000 1.1195 0.01629 0.00944 -0.1034 0.4948 1.0000
6.250 1.1315 0.01645 0.00961 -0.1004 0.4844 1.0000
6.500 1.1462 0.01665 0.00976 -0.0978 0.4747 1.0000
6.750 1.1567 0.01685 0.00993 -0.0945 0.4642 1.0000
7.000 1.1681 0.01712 0.01018 -0.0914 0.4536 1.0000
7.250 1.1823 0.01738 0.01035 -0.0889 0.4432 1.0000
7.500 1.1917 0.01773 0.01072 -0.0857 0.4318 1.0000
7.750 1.2061 0.01808 0.01098 -0.0834 0.4217 1.0000
8.000 1.2164 0.01848 0.01137 -0.0804 0.4103 1.0000
8.250 1.2281 0.01893 0.01177 -0.0779 0.3996 1.0000
8.500 1.2405 0.01940 0.01214 -0.0754 0.3891 1.0000
8.750 1.2502 0.01995 0.01270 -0.0727 0.3782 1.0000
9.000 1.2637 0.02053 0.01316 -0.0707 0.3687 1.0000
9.250 1.2746 0.02114 0.01379 -0.0683 0.3592 1.0000
9.500 1.2896 0.02178 0.01433 -0.0666 0.3508 1.0000
9.750 1.3002 0.02247 0.01505 -0.0644 0.3422 1.0000
10.000 1.3165 0.02315 0.01560 -0.0630 0.3341 1.0000
10.250 1.3254 0.02392 0.01646 -0.0607 0.3262 1.0000
10.500 1.3405 0.02464 0.01709 -0.0592 0.3189 1.0000
10.750 1.3511 0.02547 0.01796 -0.0573 0.3116 1.0000
11.000 1.3619 0.02629 0.01880 -0.0554 0.3045 1.0000
11.250 1.3781 0.02706 0.01949 -0.0542 0.2979 1.0000
11.500 1.3845 0.02805 0.02059 -0.0520 0.2912 1.0000
11.750 1.3955 0.02895 0.02146 -0.0504 0.2848 1.0000
12.000 1.4069 0.02990 0.02242 -0.0489 0.2786 1.0000
12.250 1.4138 0.03099 0.02361 -0.0470 0.2726 1.0000
12.500 1.4239 0.03199 0.02456 -0.0455 0.2664 1.0000
12.750 1.4309 0.03319 0.02582 -0.0438 0.2606 1.0000
13.000 1.4367 0.03446 0.02718 -0.0421 0.2548 1.0000
13.250 1.4464 0.03557 0.02825 -0.0408 0.2496 1.0000
13.500 1.4532 0.03691 0.02967 -0.0394 0.2443 1.0000
13.750 1.4571 0.03844 0.03131 -0.0379 0.2386 1.0000
14.000 1.4639 0.03982 0.03266 -0.0367 0.2334 1.0000
14.250 1.4690 0.04141 0.03432 -0.0355 0.2281 1.0000
14.500 1.4709 0.04328 0.03631 -0.0343 0.2223 1.0000
14.750 1.4759 0.04494 0.03794 -0.0333 0.2169 1.0000
15.000 1.4778 0.04699 0.04009 -0.0323 0.2113 1.0000
15.250 1.4782 0.04922 0.04240 -0.0315 0.2053 1.0000
15.500 1.4815 0.05116 0.04429 -0.0307 0.1995 1.0000
15.750 1.4782 0.05398 0.04728 -0.0302 0.1927 1.0000
16.000 1.4773 0.05650 0.04976 -0.0296 0.1862 1.0000
16.250 1.4730 0.05958 0.05297 -0.0293 0.1788 1.0000
16.500 1.4688 0.06257 0.05589 -0.0290 0.1714 1.0000
16.750 1.4606 0.06628 0.05974 -0.0290 0.1627 1.0000
17.000 1.4542 0.06972 0.06313 -0.0289 0.1549 1.0000
17.250 1.4463 0.07349 0.06694 -0.0291 0.1469 1.0000
17.500 1.4399 0.07710 0.07057 -0.0292 0.1398 1.0000
17.750 1.4345 0.08065 0.07411 -0.0295 0.1337 1.0000
18.000 1.4313 0.08396 0.07746 -0.0297 0.1284 1.0000
18.250 1.4282 0.08735 0.08089 -0.0302 0.1237 1.0000
18.500 1.4296 0.08995 0.08339 -0.0303 0.1200 1.0000
18.750 1.4275 0.09333 0.08691 -0.0308 0.1164 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 646 AIRFOIL (goe646-il)