GOE 443 AIRFOIL (goe443-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 443 AIRFOIL (goe443-il) Reynolds number: 50,000 Max Cl/Cd: 16.03 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe443-il-50000-n5.txt Download as CSV file: xf-goe443-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 443 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6597 0.11082 0.10412 0.0150 1.0000 0.0966
-8.750 -0.6641 0.10744 0.10083 0.0118 1.0000 0.1015
-8.500 -0.6852 0.10506 0.09861 0.0037 1.0000 0.1035
-8.250 -0.6630 0.09909 0.09261 0.0086 1.0000 0.1076
-8.000 -0.6619 0.09510 0.08860 0.0071 1.0000 0.1109
-7.750 -0.6785 0.09168 0.08533 -0.0006 1.0000 0.1184
-7.500 -0.5756 0.07178 0.06564 -0.0062 1.0000 0.0547
-7.000 -0.6440 0.07421 0.06758 -0.0075 1.0000 0.0557
-6.750 -0.6344 0.06877 0.06196 -0.0095 1.0000 0.0426
-6.500 -0.6230 0.06399 0.05686 -0.0111 1.0000 0.0363
-6.250 -0.6124 0.05956 0.05220 -0.0115 1.0000 0.0342
-6.000 -0.5997 0.05511 0.04747 -0.0119 1.0000 0.0323
-5.750 -0.5843 0.05085 0.04280 -0.0119 1.0000 0.0307
-5.500 -0.5671 0.04692 0.03840 -0.0115 1.0000 0.0301
-5.250 -0.5495 0.04342 0.03454 -0.0109 1.0000 0.0312
-5.000 -0.5297 0.04006 0.03074 -0.0102 1.0000 0.0321
-4.750 -0.5080 0.03682 0.02699 -0.0092 1.0000 0.0323
-4.500 -0.4846 0.03373 0.02337 -0.0081 1.0000 0.0320
-4.250 -0.4597 0.03087 0.02002 -0.0069 1.0000 0.0323
-4.000 -0.4337 0.02830 0.01701 -0.0057 1.0000 0.0332
-3.750 -0.4068 0.02630 0.01446 -0.0045 1.0000 0.0374
-3.500 -0.3813 0.02408 0.01213 -0.0034 1.0000 0.0426
-3.250 -0.3555 0.02240 0.01012 -0.0020 1.0000 0.0486
-3.000 -0.3313 0.02067 0.00825 -0.0012 1.0000 0.0644
-2.750 -0.2221 0.01521 0.00593 -0.0121 1.0000 1.0000
-2.500 -0.2034 0.01491 0.00518 -0.0110 1.0000 1.0000
-2.250 -0.1842 0.01466 0.00457 -0.0099 1.0000 1.0000
-2.000 -0.1645 0.01446 0.00404 -0.0088 1.0000 1.0000
-1.750 -0.1444 0.01429 0.00360 -0.0078 1.0000 1.0000
-1.500 -0.1240 0.01416 0.00317 -0.0067 1.0000 1.0000
-1.250 -0.1034 0.01406 0.00288 -0.0056 1.0000 1.0000
-1.000 -0.0828 0.01398 0.00265 -0.0045 1.0000 1.0000
-0.750 -0.0621 0.01392 0.00248 -0.0034 1.0000 1.0000
-0.500 -0.0414 0.01388 0.00234 -0.0023 1.0000 1.0000
-0.250 -0.0207 0.01386 0.00226 -0.0011 1.0000 1.0000
0.000 0.0000 0.01385 0.00223 0.0000 1.0000 1.0000
0.250 0.0207 0.01386 0.00226 0.0011 1.0000 1.0000
0.500 0.0414 0.01388 0.00234 0.0023 1.0000 1.0000
0.750 0.0621 0.01392 0.00248 0.0034 1.0000 1.0000
1.000 0.0827 0.01398 0.00265 0.0045 1.0000 1.0000
1.250 0.1034 0.01406 0.00288 0.0056 1.0000 1.0000
1.500 0.1239 0.01417 0.00317 0.0067 1.0000 1.0000
1.750 0.1442 0.01429 0.00361 0.0078 1.0000 1.0000
2.000 0.1643 0.01446 0.00405 0.0089 1.0000 1.0000
2.250 0.1840 0.01466 0.00457 0.0100 1.0000 1.0000
2.500 0.2032 0.01491 0.00519 0.0110 1.0000 1.0000
2.750 0.2219 0.01521 0.00594 0.0122 1.0000 1.0000
3.000 0.3313 0.02067 0.00825 0.0012 0.0644 1.0000
3.250 0.3555 0.02239 0.01012 0.0020 0.0486 1.0000
3.500 0.3814 0.02408 0.01213 0.0034 0.0426 1.0000
3.750 0.4068 0.02631 0.01447 0.0045 0.0374 1.0000
4.000 0.4338 0.02831 0.01701 0.0057 0.0332 1.0000
4.250 0.4598 0.03088 0.02002 0.0069 0.0323 1.0000
4.500 0.4846 0.03372 0.02336 0.0081 0.0321 1.0000
4.750 0.5080 0.03681 0.02698 0.0092 0.0324 1.0000
5.000 0.5297 0.04006 0.03075 0.0102 0.0321 1.0000
5.250 0.5495 0.04341 0.03455 0.0110 0.0314 1.0000
5.500 0.5672 0.04692 0.03842 0.0115 0.0304 1.0000
5.750 0.5841 0.05084 0.04278 0.0119 0.0307 1.0000
6.000 0.5994 0.05511 0.04746 0.0119 0.0323 1.0000
6.250 0.6122 0.05951 0.05216 0.0115 0.0342 1.0000
6.500 0.6226 0.06399 0.05687 0.0112 0.0363 1.0000
6.750 0.6341 0.06878 0.06197 0.0096 0.0426 1.0000
7.000 0.6438 0.07410 0.06747 0.0076 0.0555 1.0000
7.500 0.5751 0.07157 0.06542 0.0064 0.0545 1.0000
8.000 0.5639 0.08420 0.07810 -0.0033 0.1250 1.0000
8.250 0.5636 0.08846 0.08233 -0.0041 0.1213 1.0000
8.500 0.5794 0.09407 0.08793 -0.0017 0.1174 1.0000
8.750 0.5543 0.09685 0.09059 -0.0071 0.1119 1.0000
9.000 0.5542 0.10092 0.09463 -0.0081 0.1082 1.0000
9.250 0.5689 0.10611 0.09992 -0.0063 0.1044 1.0000
9.500 0.5655 0.11046 0.10419 -0.0080 0.1031 1.0000
9.750 0.5501 0.11308 0.10666 -0.0118 0.0987 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 443 AIRFOIL (goe443-il)