ONERA OA206 AIRFOIL (oa206-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA OA206 AIRFOIL (oa206-il) Reynolds number: 50,000 Max Cl/Cd: 27.32 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oa206-il-50000.txt Download as CSV file: xf-oa206-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: ONERA OA206 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.6583 0.09753 0.09138 0.0351 1.0000 0.2665
-7.250 -0.6510 0.09372 0.08757 0.0364 1.0000 0.2897
-7.000 -0.6611 0.09096 0.08491 0.0336 1.0000 0.3116
-6.750 -0.6389 0.08688 0.08082 0.0398 1.0000 0.3469
-6.500 -0.6402 0.08408 0.07809 0.0408 1.0000 0.3819
-6.250 -0.6345 0.08121 0.07528 0.0445 1.0000 0.4254
-6.000 -0.5973 0.07726 0.07126 0.0529 1.0000 0.4915
-5.500 -0.5118 0.06990 0.06371 0.0670 1.0000 0.7065
-4.500 -0.4689 0.04065 0.03191 -0.0105 1.0000 0.1686
-4.250 -0.4378 0.03685 0.02756 -0.0111 1.0000 0.1507
-4.000 -0.4061 0.03393 0.02385 -0.0113 1.0000 0.1406
-3.750 -0.3760 0.03095 0.02043 -0.0112 1.0000 0.1333
-3.500 -0.3443 0.02884 0.01764 -0.0107 1.0000 0.1277
-3.250 -0.3148 0.02660 0.01510 -0.0102 1.0000 0.1270
-3.000 -0.2862 0.02457 0.01294 -0.0097 1.0000 0.1324
-2.750 -0.2574 0.02303 0.01123 -0.0090 1.0000 0.1416
-2.500 -0.2287 0.02137 0.00955 -0.0080 1.0000 0.1499
-2.250 -0.2021 0.02001 0.00819 -0.0069 1.0000 0.1674
-2.000 -0.1309 0.01495 0.00634 -0.0093 1.0000 1.0000
-1.750 -0.1068 0.01481 0.00557 -0.0084 1.0000 1.0000
-1.500 -0.0841 0.01469 0.00506 -0.0077 1.0000 1.0000
-1.250 -0.0619 0.01459 0.00471 -0.0069 1.0000 1.0000
-1.000 -0.0403 0.01451 0.00443 -0.0061 1.0000 1.0000
-0.750 -0.0193 0.01446 0.00422 -0.0052 1.0000 1.0000
-0.500 0.0010 0.01444 0.00407 -0.0041 1.0000 1.0000
-0.250 0.0208 0.01444 0.00395 -0.0030 1.0000 1.0000
0.000 0.0401 0.01446 0.00390 -0.0018 1.0000 1.0000
0.250 0.0590 0.01451 0.00389 -0.0005 1.0000 1.0000
0.500 0.0780 0.01459 0.00393 0.0007 1.0000 1.0000
0.750 0.0974 0.01469 0.00402 0.0017 1.0000 1.0000
1.000 0.1172 0.01483 0.00416 0.0026 1.0000 1.0000
1.250 0.1375 0.01500 0.00435 0.0034 1.0000 1.0000
1.500 0.1581 0.01521 0.00460 0.0040 1.0000 1.0000
1.750 0.1791 0.01546 0.00492 0.0045 1.0000 1.0000
2.000 0.2002 0.01577 0.00532 0.0047 1.0000 1.0000
2.250 0.2213 0.01614 0.00584 0.0048 1.0000 1.0000
2.500 0.2420 0.01661 0.00645 0.0046 1.0000 1.0000
2.750 0.2620 0.01723 0.00724 0.0041 1.0000 1.0000
3.000 0.3423 0.01820 0.00874 -0.0079 0.9654 1.0000
3.250 0.4470 0.01817 0.00945 -0.0190 0.8674 1.0000
3.500 0.4694 0.01786 0.00924 -0.0137 0.7830 1.0000
3.750 0.4831 0.01783 0.00903 -0.0073 0.6988 1.0000
4.000 0.4991 0.01827 0.00918 -0.0025 0.6149 1.0000
4.250 0.5186 0.01912 0.00963 0.0008 0.5380 1.0000
4.500 0.5402 0.02025 0.01042 0.0030 0.4684 1.0000
4.750 0.5631 0.02160 0.01154 0.0046 0.4023 1.0000
5.000 0.5853 0.02300 0.01270 0.0060 0.3338 1.0000
5.250 0.6067 0.02438 0.01374 0.0074 0.2570 1.0000
5.500 0.6295 0.02637 0.01535 0.0085 0.1873 1.0000
5.750 0.6552 0.02881 0.01788 0.0093 0.1503 1.0000
6.000 0.6805 0.03136 0.02052 0.0098 0.1294 1.0000
6.250 0.7059 0.03436 0.02369 0.0102 0.1178 1.0000
6.500 0.7309 0.03831 0.02851 0.0105 0.1130 1.0000
6.750 0.7532 0.04151 0.03175 0.0107 0.1050 1.0000
7.000 0.7723 0.04613 0.03725 0.0104 0.1028 1.0000
7.250 0.7891 0.05138 0.04301 0.0100 0.1038 1.0000
7.500 0.7951 0.05905 0.05166 0.0072 0.1102 1.0000
7.750 0.8014 0.06581 0.05875 0.0050 0.1158 1.0000
8.000 0.7949 0.07503 0.06837 -0.0011 0.1284 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA OA206 AIRFOIL (oa206-il)