NACA 16-006 (naca16006-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 16-006 (naca16006-il) Reynolds number: 500,000 Max Cl/Cd: 41.23 at α=1.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca16006-il-500000.txt Download as CSV file: xf-naca16006-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 16-006
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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5229 0.08508 0.08289 -0.0131 1.0000 0.0126
-8.750 -0.5268 0.08032 0.07814 -0.0151 1.0000 0.0127
-8.500 -0.5327 0.07530 0.07314 -0.0179 1.0000 0.0127
-8.250 -0.5430 0.06995 0.06779 -0.0222 1.0000 0.0127
-8.000 -0.5547 0.06558 0.06339 -0.0236 1.0000 0.0126
-7.750 -0.5646 0.06158 0.05934 -0.0233 1.0000 0.0127
-7.500 -0.5696 0.05726 0.05496 -0.0230 1.0000 0.0127
-7.250 -0.5726 0.05305 0.05066 -0.0220 1.0000 0.0127
-7.000 -0.5744 0.04895 0.04645 -0.0204 1.0000 0.0127
-6.750 -0.5745 0.04499 0.04238 -0.0185 1.0000 0.0128
-6.500 -0.5737 0.04109 0.03834 -0.0161 1.0000 0.0128
-6.250 -0.5719 0.03734 0.03444 -0.0135 1.0000 0.0128
-6.000 -0.5688 0.03379 0.03072 -0.0106 1.0000 0.0128
-5.750 -0.5648 0.03037 0.02711 -0.0076 1.0000 0.0129
-5.500 -0.5596 0.02706 0.02358 -0.0045 1.0000 0.0129
-5.250 -0.5684 0.01976 0.01585 0.0001 1.0000 0.0136
-5.000 -0.5617 0.01643 0.01227 0.0029 1.0000 0.0144
-4.750 -0.5493 0.01438 0.01005 0.0051 1.0000 0.0153
-4.500 -0.5345 0.01275 0.00825 0.0071 1.0000 0.0172
-4.250 -0.5097 0.01451 0.00988 0.0088 1.0000 0.0230
-3.500 -0.4567 0.01478 0.00851 0.0152 1.0000 0.0142
-3.250 -0.4319 0.01337 0.00691 0.0162 1.0000 0.0139
-3.000 -0.4082 0.01182 0.00525 0.0172 1.0000 0.0153
-2.750 -0.3844 0.01119 0.00458 0.0180 1.0000 0.0180
-2.500 -0.3603 0.01080 0.00414 0.0187 1.0000 0.0213
-2.250 -0.3362 0.01019 0.00344 0.0194 1.0000 0.0253
-2.000 -0.3117 0.00972 0.00293 0.0200 1.0000 0.0356
-1.750 -0.2870 0.00929 0.00255 0.0205 1.0000 0.0592
-1.500 -0.2642 0.00644 0.00217 0.0200 1.0000 0.7026
-1.250 -0.2461 0.00602 0.00254 0.0232 1.0000 0.9237
-1.000 -0.2158 0.00623 0.00273 0.0231 1.0000 0.9625
-0.750 -0.1513 0.00658 0.00301 0.0154 1.0000 0.9859
-0.250 -0.0074 0.00684 0.00316 -0.0037 1.0000 1.0000
0.000 0.0000 0.00684 0.00315 0.0000 1.0000 1.0000
0.250 0.0074 0.00684 0.00316 0.0037 1.0000 1.0000
0.750 0.1492 0.00659 0.00301 -0.0150 0.9862 1.0000
1.000 0.2156 0.00623 0.00273 -0.0230 0.9625 1.0000
1.250 0.2458 0.00602 0.00253 -0.0232 0.9242 1.0000
1.500 0.2639 0.00640 0.00217 -0.0200 0.7106 1.0000
1.750 0.2867 0.00929 0.00255 -0.0204 0.0593 1.0000
2.000 0.3113 0.00972 0.00292 -0.0199 0.0356 1.0000
2.250 0.3359 0.01017 0.00342 -0.0193 0.0254 1.0000
2.500 0.3600 0.01079 0.00413 -0.0186 0.0213 1.0000
2.750 0.3840 0.01119 0.00457 -0.0179 0.0180 1.0000
3.000 0.4078 0.01181 0.00524 -0.0171 0.0153 1.0000
3.250 0.4315 0.01337 0.00691 -0.0161 0.0138 1.0000
3.500 0.4562 0.01476 0.00850 -0.0151 0.0142 1.0000
4.250 0.5259 0.02324 0.01815 -0.0094 0.0205 1.0000
4.500 0.5456 0.02476 0.01991 -0.0074 0.0175 1.0000
4.750 0.5623 0.02639 0.02172 -0.0055 0.0154 1.0000
5.000 0.5765 0.02853 0.02402 -0.0035 0.0143 1.0000
5.250 0.5847 0.03240 0.02818 -0.0006 0.0135 1.0000
5.500 0.5815 0.03947 0.03569 0.0038 0.0129 1.0000
5.750 0.5900 0.04274 0.03920 0.0067 0.0128 1.0000
6.000 0.5976 0.04608 0.04275 0.0095 0.0128 1.0000
6.250 0.6046 0.04950 0.04635 0.0121 0.0128 1.0000
6.500 0.6112 0.05300 0.05002 0.0144 0.0128 1.0000
6.750 0.6174 0.05659 0.05377 0.0163 0.0127 1.0000
7.000 0.6231 0.06026 0.05757 0.0177 0.0127 1.0000
7.250 0.6290 0.06399 0.06142 0.0186 0.0126 1.0000
7.500 0.6357 0.06774 0.06532 0.0187 0.0124 1.0000
7.750 0.6442 0.07148 0.06915 0.0178 0.0120 1.0000
8.000 0.6486 0.07591 0.07365 0.0158 0.0116 1.0000
8.250 0.6477 0.08052 0.07831 0.0132 0.0114 1.0000
8.500 0.6420 0.08529 0.08309 0.0094 0.0114 1.0000
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Polar data table (+)
Polar graphs
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