NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA 63A010 AIRFOIL (n63010a-il) Reynolds number: 500,000 Max Cl/Cd: 52.34 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n63010a-il-500000.txt Download as CSV file: xf-n63010a-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 63A010 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.7715 0.06804 0.06570 -0.0204 1.0000 0.0217
-9.750 -0.7931 0.06297 0.06050 -0.0221 1.0000 0.0217
-9.500 -0.8092 0.05877 0.05613 -0.0214 1.0000 0.0220
-9.250 -0.8065 0.05723 0.05420 -0.0200 1.0000 0.0243
-9.000 -0.8095 0.05408 0.05078 -0.0186 1.0000 0.0244
-8.500 -0.8409 0.03462 0.03002 -0.0128 1.0000 0.0181
-8.250 -0.8310 0.02970 0.02482 -0.0111 1.0000 0.0170
-8.000 -0.8160 0.02668 0.02145 -0.0096 1.0000 0.0170
-7.750 -0.7977 0.02438 0.01885 -0.0082 1.0000 0.0172
-7.500 -0.7773 0.02265 0.01688 -0.0071 1.0000 0.0174
-7.250 -0.7563 0.02119 0.01517 -0.0061 1.0000 0.0180
-7.000 -0.7370 0.01854 0.01234 -0.0049 1.0000 0.0189
-6.750 -0.7156 0.01728 0.01102 -0.0039 1.0000 0.0194
-6.500 -0.6941 0.01633 0.01002 -0.0028 1.0000 0.0201
-6.250 -0.6729 0.01550 0.00914 -0.0017 1.0000 0.0208
-6.000 -0.6521 0.01473 0.00830 -0.0005 1.0000 0.0216
-5.750 -0.6317 0.01404 0.00755 0.0009 1.0000 0.0225
-5.500 -0.6106 0.01354 0.00701 0.0021 1.0000 0.0237
-5.250 -0.5893 0.01316 0.00660 0.0033 1.0000 0.0245
-5.000 -0.5738 0.01202 0.00539 0.0054 1.0000 0.0265
-4.750 -0.5544 0.01153 0.00487 0.0068 1.0000 0.0288
-4.500 -0.5255 0.01112 0.00442 0.0063 0.9982 0.0322
-4.250 -0.4884 0.01050 0.00381 0.0041 0.9945 0.0426
-4.000 -0.4537 0.00977 0.00329 0.0022 0.9893 0.0846
-3.750 -0.4201 0.00860 0.00278 -0.0001 0.9845 0.2312
-3.500 -0.3895 0.00744 0.00239 -0.0016 0.9776 0.4153
-3.250 -0.3564 0.00689 0.00224 -0.0031 0.9708 0.5216
-3.000 -0.3259 0.00659 0.00213 -0.0037 0.9606 0.5831
-2.750 -0.2946 0.00643 0.00204 -0.0044 0.9505 0.6149
-2.500 -0.2651 0.00631 0.00197 -0.0046 0.9389 0.6427
-2.250 -0.2379 0.00621 0.00191 -0.0042 0.9258 0.6714
-2.000 -0.2120 0.00615 0.00187 -0.0036 0.9118 0.6979
-1.750 -0.1860 0.00611 0.00183 -0.0029 0.8978 0.7180
-1.500 -0.1599 0.00608 0.00179 -0.0024 0.8839 0.7348
-1.250 -0.1336 0.00606 0.00175 -0.0019 0.8705 0.7497
-1.000 -0.1071 0.00604 0.00172 -0.0014 0.8574 0.7628
-0.750 -0.0805 0.00603 0.00169 -0.0010 0.8448 0.7742
-0.500 -0.0537 0.00602 0.00167 -0.0007 0.8326 0.7856
-0.250 -0.0268 0.00602 0.00165 -0.0003 0.8207 0.7973
0.000 0.0000 0.00603 0.00164 0.0000 0.8092 0.8092
0.250 0.0268 0.00602 0.00165 0.0004 0.7973 0.8207
0.500 0.0536 0.00602 0.00167 0.0007 0.7857 0.8326
0.750 0.0804 0.00603 0.00169 0.0010 0.7743 0.8448
1.000 0.1071 0.00604 0.00172 0.0014 0.7628 0.8574
1.250 0.1336 0.00606 0.00175 0.0019 0.7497 0.8705
1.500 0.1599 0.00608 0.00179 0.0024 0.7348 0.8839
1.750 0.1860 0.00611 0.00183 0.0030 0.7180 0.8978
2.000 0.2119 0.00615 0.00187 0.0036 0.6978 0.9118
2.250 0.2379 0.00621 0.00191 0.0042 0.6714 0.9258
2.500 0.2650 0.00631 0.00196 0.0046 0.6423 0.9389
2.750 0.2945 0.00643 0.00204 0.0044 0.6149 0.9505
3.000 0.3258 0.00659 0.00213 0.0037 0.5832 0.9606
3.250 0.3564 0.00688 0.00224 0.0031 0.5221 0.9708
3.500 0.3894 0.00744 0.00239 0.0016 0.4152 0.9776
3.750 0.4199 0.00859 0.00278 0.0001 0.2318 0.9845
4.000 0.4536 0.00976 0.00328 -0.0021 0.0847 0.9893
4.250 0.4884 0.01050 0.00381 -0.0041 0.0427 0.9945
4.500 0.5254 0.01112 0.00442 -0.0063 0.0322 0.9982
4.750 0.5541 0.01153 0.00487 -0.0068 0.0288 1.0000
5.000 0.5736 0.01202 0.00538 -0.0054 0.0266 1.0000
5.250 0.5890 0.01316 0.00659 -0.0032 0.0245 1.0000
5.500 0.6104 0.01353 0.00701 -0.0020 0.0237 1.0000
5.750 0.6315 0.01403 0.00754 -0.0008 0.0225 1.0000
6.000 0.6520 0.01473 0.00830 0.0005 0.0216 1.0000
6.250 0.6727 0.01550 0.00913 0.0017 0.0208 1.0000
6.500 0.6940 0.01633 0.01002 0.0029 0.0201 1.0000
6.750 0.7156 0.01728 0.01102 0.0039 0.0194 1.0000
7.000 0.7370 0.01853 0.01234 0.0049 0.0189 1.0000
7.250 0.7562 0.02123 0.01521 0.0061 0.0180 1.0000
7.500 0.7773 0.02265 0.01688 0.0071 0.0174 1.0000
7.750 0.7978 0.02439 0.01886 0.0082 0.0172 1.0000
8.000 0.8160 0.02670 0.02146 0.0095 0.0170 1.0000
8.250 0.8310 0.02972 0.02484 0.0111 0.0170 1.0000
8.500 0.8412 0.03459 0.02998 0.0127 0.0181 1.0000
14.750 0.5255 0.15208 0.14991 -0.0125 0.0178 1.0000
15.000 0.5253 0.15525 0.15307 -0.0139 0.0178 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 63A010 AIRFOIL (n63010a-il)