FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il) Reynolds number: 100,000 Max Cl/Cd: 13.26 at α=1° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx66s171-il-100000.txt Download as CSV file: xf-fx66s171-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-S-171 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2561 0.09573 0.09186 -0.0722 0.9338 0.1594
-9.750 -0.1960 0.09131 0.08722 -0.0740 0.9142 0.1664
-9.500 -0.3680 0.05575 0.05108 -0.1096 0.8928 0.0702
-9.250 -0.3734 0.05117 0.04618 -0.1112 0.8724 0.0675
-9.000 -0.3869 0.04655 0.04097 -0.1112 0.8555 0.0638
-8.750 -0.3941 0.04350 0.03698 -0.1097 0.8411 0.0607
-8.500 -0.3808 0.04123 0.03426 -0.1089 0.8296 0.0603
-8.250 -0.3669 0.03823 0.03084 -0.1081 0.8194 0.0604
-8.000 -0.3514 0.03505 0.02750 -0.1077 0.8091 0.0626
-7.750 -0.3309 0.03342 0.02571 -0.1071 0.8003 0.0656
-7.500 -0.3111 0.03176 0.02375 -0.1062 0.7912 0.0684
-7.250 -0.2893 0.03018 0.02181 -0.1052 0.7833 0.0712
-7.000 -0.2667 0.02846 0.01999 -0.1046 0.7758 0.0761
-6.750 -0.2434 0.02735 0.01879 -0.1039 0.7694 0.0825
-6.500 -0.2209 0.02611 0.01753 -0.1031 0.7624 0.0896
-6.250 -0.1972 0.02526 0.01655 -0.1022 0.7572 0.0999
-6.000 -0.1785 0.02430 0.01579 -0.1010 0.7504 0.1133
-5.750 -0.1608 0.02324 0.01492 -0.0997 0.7447 0.1343
-5.500 -0.1462 0.02175 0.01381 -0.0984 0.7399 0.1906
-5.250 -0.1373 0.02086 0.01420 -0.0962 0.7338 0.4123
-5.000 -0.1128 0.02199 0.01519 -0.0946 0.7289 0.4695
-4.750 -0.0874 0.02297 0.01603 -0.0931 0.7250 0.5000
-4.500 -0.0658 0.02375 0.01675 -0.0917 0.7192 0.5229
-4.250 -0.0416 0.02453 0.01751 -0.0899 0.7145 0.5421
-4.000 -0.0164 0.02517 0.01805 -0.0882 0.7108 0.5612
-3.750 0.0040 0.02581 0.01869 -0.0866 0.7059 0.5774
-3.500 0.0258 0.02630 0.01914 -0.0851 0.7013 0.5928
-3.250 0.0500 0.02654 0.01927 -0.0839 0.6975 0.6068
-3.000 0.0752 0.02664 0.01922 -0.0832 0.6943 0.6193
-2.750 0.0936 0.02698 0.01953 -0.0825 0.6892 0.6311
-2.500 0.1154 0.02722 0.01973 -0.0815 0.6852 0.6397
-2.250 0.1412 0.02716 0.01953 -0.0815 0.6819 0.6489
-2.000 0.1689 0.02711 0.01932 -0.0815 0.6792 0.6577
-1.750 0.1840 0.02765 0.01989 -0.0807 0.6738 0.6646
-1.500 0.2054 0.02795 0.02014 -0.0804 0.6699 0.6722
-1.250 0.2309 0.02803 0.02012 -0.0806 0.6669 0.6796
-1.000 0.2592 0.02803 0.01999 -0.0807 0.6645 0.6865
-0.750 0.2729 0.02887 0.02085 -0.0801 0.6594 0.6936
-0.500 0.2897 0.02947 0.02147 -0.0792 0.6548 0.7002
-0.250 0.3147 0.02975 0.02165 -0.0795 0.6518 0.7084
0.000 0.3426 0.02980 0.02165 -0.0794 0.6494 0.7152
0.250 0.3471 0.03135 0.02325 -0.0781 0.6435 0.7225
0.500 0.3610 0.03215 0.02408 -0.0770 0.6385 0.7288
0.750 0.3908 0.03219 0.02404 -0.0774 0.6355 0.7376
1.000 0.4238 0.03196 0.02376 -0.0776 0.6334 0.7456
1.250 0.3972 0.03533 0.02728 -0.0738 0.6239 0.7532
1.500 0.4231 0.03550 0.02746 -0.0735 0.6207 0.7621
1.750 0.4592 0.03522 0.02713 -0.0742 0.6187 0.7726
2.000 0.3837 0.04120 0.03331 -0.0666 0.6080 0.7790
2.250 0.4197 0.04102 0.03312 -0.0673 0.6051 0.7902
2.500 0.4680 0.04030 0.03237 -0.0690 0.6033 0.8029
2.750 0.3886 0.04669 0.03889 -0.0621 0.5938 0.8126
3.000 0.4109 0.04741 0.03964 -0.0620 0.5904 0.8278
3.250 0.4550 0.04683 0.03911 -0.0629 0.5880 0.8477
4.000 0.4943 0.05245 0.04501 -0.0657 0.5731 1.0000
4.250 0.4820 0.05598 0.04849 -0.0657 0.5670 1.0000
4.500 0.4821 0.05884 0.05129 -0.0661 0.5640 1.0000
4.750 0.5099 0.05984 0.05222 -0.0671 0.5596 1.0000
5.000 0.4095 0.07105 0.06359 -0.0671 0.6099 1.0000
5.250 0.4425 0.07255 0.06503 -0.0684 0.6045 1.0000
5.500 0.4880 0.07485 0.06726 -0.0707 0.6021 1.0000
5.750 0.4576 0.07619 0.06860 -0.0678 0.5926 1.0000
6.000 0.4925 0.07794 0.07031 -0.0691 0.5885 1.0000
6.250 0.4852 0.07984 0.07221 -0.0679 0.5807 1.0000
6.500 0.5074 0.08148 0.07384 -0.0683 0.5752 1.0000
6.750 0.5477 0.08361 0.07596 -0.0699 0.5722 1.0000
7.000 0.5288 0.08523 0.07759 -0.0680 0.5621 1.0000
7.250 0.5605 0.08708 0.07944 -0.0689 0.5581 1.0000
7.500 0.5546 0.08927 0.08165 -0.0681 0.5510 1.0000
7.750 0.5783 0.09085 0.08326 -0.0685 0.5442 1.0000
8.000 0.6195 0.09320 0.08564 -0.0699 0.5415 1.0000
8.250 0.5949 0.09501 0.08747 -0.0683 0.5314 1.0000
8.500 0.6335 0.09694 0.08945 -0.0693 0.5272 1.0000
8.750 0.6165 0.09930 0.09184 -0.0684 0.5189 1.0000
9.000 0.6459 0.10106 0.09365 -0.0689 0.5133 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-S-171 AIRFOIL (fx66s171-il)