NACA 64-008A AIRFOIL (n64008a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 64-008A AIRFOIL (n64008a-il) Reynolds number: 50,000 Max Cl/Cd: 19.06 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64008a-il-50000.txt Download as CSV file: xf-n64008a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-008A 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
------ -------- --------- --------- -------- -------- --------
-6.500 -0.6776 0.05318 0.04526 -0.0137 1.0000 0.1538
-6.250 -0.6581 0.04768 0.03931 -0.0137 1.0000 0.1334
-6.000 -0.6385 0.04335 0.03430 -0.0129 1.0000 0.1207
-5.750 -0.6169 0.03981 0.02991 -0.0117 1.0000 0.1124
-5.500 -0.5944 0.03658 0.02617 -0.0107 1.0000 0.1114
-5.250 -0.5707 0.03411 0.02306 -0.0094 1.0000 0.1138
-5.000 -0.5447 0.03118 0.01984 -0.0085 1.0000 0.1154
-4.750 -0.5167 0.02860 0.01707 -0.0078 1.0000 0.1177
-4.500 -0.4873 0.02645 0.01478 -0.0069 1.0000 0.1233
-4.250 -0.4593 0.02459 0.01290 -0.0060 1.0000 0.1384
-4.000 -0.4333 0.02272 0.01108 -0.0048 1.0000 0.1595
-3.750 -0.2491 0.02188 0.01257 -0.0098 1.0000 0.9931
-3.500 -0.2151 0.02087 0.01115 -0.0131 1.0000 1.0000
-3.250 -0.1968 0.02014 0.01020 -0.0134 1.0000 1.0000
-3.000 -0.1783 0.01948 0.00935 -0.0135 1.0000 1.0000
-2.750 -0.1598 0.01888 0.00859 -0.0135 1.0000 1.0000
-2.500 -0.1413 0.01834 0.00791 -0.0134 1.0000 1.0000
-2.250 -0.1231 0.01785 0.00731 -0.0132 1.0000 1.0000
-2.000 -0.1050 0.01742 0.00674 -0.0128 1.0000 1.0000
-1.750 -0.0874 0.01703 0.00628 -0.0123 1.0000 1.0000
-1.500 -0.0704 0.01670 0.00589 -0.0115 1.0000 1.0000
-1.250 -0.0543 0.01641 0.00557 -0.0105 1.0000 1.0000
-1.000 -0.0395 0.01616 0.00531 -0.0093 1.0000 1.0000
-0.750 -0.0264 0.01597 0.00509 -0.0076 1.0000 1.0000
-0.500 -0.0155 0.01582 0.00495 -0.0055 1.0000 1.0000
-0.250 -0.0072 0.01573 0.00488 -0.0028 1.0000 1.0000
0.000 0.0000 0.01570 0.00485 0.0000 1.0000 1.0000
0.250 0.0072 0.01573 0.00488 0.0028 1.0000 1.0000
0.500 0.0156 0.01582 0.00495 0.0054 1.0000 1.0000
0.750 0.0264 0.01596 0.00509 0.0076 1.0000 1.0000
1.000 0.0395 0.01616 0.00530 0.0093 1.0000 1.0000
1.250 0.0544 0.01640 0.00556 0.0105 1.0000 1.0000
1.500 0.0705 0.01669 0.00589 0.0115 1.0000 1.0000
1.750 0.0875 0.01703 0.00628 0.0123 1.0000 1.0000
2.000 0.1051 0.01741 0.00674 0.0128 1.0000 1.0000
2.250 0.1231 0.01785 0.00730 0.0132 1.0000 1.0000
2.500 0.1414 0.01833 0.00790 0.0134 1.0000 1.0000
2.750 0.1599 0.01887 0.00858 0.0135 1.0000 1.0000
3.000 0.1785 0.01947 0.00934 0.0135 1.0000 1.0000
3.250 0.1970 0.02013 0.01019 0.0133 1.0000 1.0000
3.500 0.2153 0.02086 0.01114 0.0131 1.0000 1.0000
3.750 0.2495 0.02187 0.01256 0.0097 0.9930 1.0000
4.000 0.4333 0.02273 0.01108 0.0048 0.1595 1.0000
4.250 0.4592 0.02459 0.01290 0.0060 0.1384 1.0000
4.500 0.4872 0.02645 0.01478 0.0069 0.1233 1.0000
4.750 0.5166 0.02860 0.01707 0.0078 0.1177 1.0000
5.000 0.5447 0.03118 0.01983 0.0085 0.1153 1.0000
5.250 0.5707 0.03411 0.02306 0.0094 0.1138 1.0000
5.500 0.5944 0.03658 0.02616 0.0107 0.1114 1.0000
5.750 0.6169 0.03982 0.02992 0.0117 0.1124 1.0000
6.000 0.6385 0.04335 0.03430 0.0129 0.1207 1.0000
6.250 0.6582 0.04769 0.03932 0.0136 0.1334 1.0000
6.500 0.6778 0.05320 0.04528 0.0137 0.1539 1.0000
8.250 0.6249 0.10021 0.09354 -0.0334 0.3816 1.0000
8.500 0.6488 0.10514 0.09850 -0.0316 0.3606 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 64-008A AIRFOIL (n64008a-il)