NACA 0006 (naca0006-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 0006 (naca0006-il) Reynolds number: 50,000 Max Cl/Cd: 24.9 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca0006-il-50000.txt Download as CSV file: xf-naca0006-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0006
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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6744 0.10463 0.09811 0.0361 1.0000 0.2841
-8.000 -0.6770 0.10152 0.09507 0.0360 1.0000 0.3021
-7.750 -0.6873 0.09891 0.09256 0.0351 1.0000 0.3183
-7.500 -0.6648 0.09391 0.08753 0.0381 1.0000 0.3430
-7.000 -0.6541 0.08634 0.08004 0.0401 1.0000 0.3826
-6.750 -0.6562 0.08314 0.07691 0.0410 1.0000 0.4060
-6.500 -0.6571 0.08012 0.07396 0.0424 1.0000 0.4327
-6.250 -0.6588 0.05709 0.05000 -0.0072 1.0000 0.1590
-6.000 -0.6399 0.04972 0.04189 -0.0102 1.0000 0.1308
-5.750 -0.6193 0.04421 0.03547 -0.0109 1.0000 0.1188
-5.500 -0.5974 0.03960 0.03037 -0.0106 1.0000 0.1141
-5.250 -0.5736 0.03565 0.02578 -0.0099 1.0000 0.1133
-5.000 -0.5484 0.03269 0.02204 -0.0090 1.0000 0.1202
-4.750 -0.5222 0.02938 0.01850 -0.0083 1.0000 0.1264
-4.500 -0.4962 0.02697 0.01584 -0.0074 1.0000 0.1435
-4.250 -0.4680 0.02457 0.01319 -0.0064 1.0000 0.1608
-4.000 -0.4405 0.02247 0.01106 -0.0054 1.0000 0.1897
-3.750 -0.4135 0.02034 0.00925 -0.0046 1.0000 0.2448
-3.500 -0.4071 0.01635 0.00795 0.0011 1.0000 0.6026
-3.250 -0.2716 0.01678 0.00780 -0.0089 1.0000 1.0000
-3.000 -0.2535 0.01614 0.00684 -0.0090 1.0000 1.0000
-2.750 -0.2342 0.01567 0.00607 -0.0088 1.0000 1.0000
-2.500 -0.2140 0.01531 0.00544 -0.0083 1.0000 1.0000
-2.250 -0.1933 0.01503 0.00489 -0.0078 1.0000 1.0000
-2.000 -0.1723 0.01481 0.00445 -0.0071 1.0000 1.0000
-1.750 -0.1511 0.01464 0.00409 -0.0063 1.0000 1.0000
-1.500 -0.1297 0.01450 0.00379 -0.0055 1.0000 1.0000
-1.250 -0.1081 0.01439 0.00352 -0.0047 1.0000 1.0000
-1.000 -0.0866 0.01431 0.00332 -0.0038 1.0000 1.0000
-0.750 -0.0650 0.01424 0.00318 -0.0029 1.0000 1.0000
-0.500 -0.0434 0.01420 0.00307 -0.0019 1.0000 1.0000
-0.250 -0.0217 0.01417 0.00300 -0.0010 1.0000 1.0000
0.000 0.0000 0.01416 0.00298 0.0000 1.0000 1.0000
0.250 0.0217 0.01417 0.00300 0.0010 1.0000 1.0000
0.500 0.0434 0.01420 0.00307 0.0019 1.0000 1.0000
0.750 0.0650 0.01424 0.00317 0.0028 1.0000 1.0000
1.000 0.0866 0.01430 0.00332 0.0038 1.0000 1.0000
1.250 0.1082 0.01439 0.00352 0.0047 1.0000 1.0000
1.500 0.1297 0.01450 0.00379 0.0055 1.0000 1.0000
1.750 0.1512 0.01464 0.00409 0.0063 1.0000 1.0000
2.000 0.1724 0.01481 0.00445 0.0071 1.0000 1.0000
2.250 0.1934 0.01503 0.00488 0.0077 1.0000 1.0000
2.500 0.2141 0.01531 0.00544 0.0083 1.0000 1.0000
2.750 0.2343 0.01567 0.00607 0.0087 1.0000 1.0000
3.000 0.2537 0.01614 0.00684 0.0090 1.0000 1.0000
3.250 0.2718 0.01678 0.00780 0.0089 1.0000 1.0000
3.500 0.4071 0.01635 0.00795 -0.0011 0.6032 1.0000
3.750 0.4135 0.02034 0.00925 0.0046 0.2449 1.0000
4.000 0.4405 0.02247 0.01106 0.0054 0.1897 1.0000
4.250 0.4680 0.02457 0.01319 0.0064 0.1608 1.0000
4.500 0.4962 0.02697 0.01583 0.0074 0.1435 1.0000
4.750 0.5222 0.02938 0.01849 0.0083 0.1264 1.0000
5.000 0.5484 0.03269 0.02204 0.0090 0.1202 1.0000
5.250 0.5736 0.03565 0.02578 0.0099 0.1133 1.0000
5.500 0.5974 0.03960 0.03037 0.0106 0.1141 1.0000
5.750 0.6194 0.04421 0.03548 0.0109 0.1188 1.0000
6.000 0.6399 0.04973 0.04189 0.0102 0.1308 1.0000
6.250 0.6589 0.05710 0.05001 0.0072 0.1590 1.0000
7.000 0.6546 0.08637 0.08007 -0.0401 0.3824 1.0000
7.500 0.6649 0.09390 0.08752 -0.0383 0.3427 1.0000
7.750 0.6872 0.09890 0.09255 -0.0353 0.3183 1.0000
8.000 0.6770 0.10150 0.09505 -0.0362 0.3020 1.0000
8.250 0.6750 0.10464 0.09813 -0.0362 0.2836 1.0000
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Polar data table (+)
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