NACA 0006 (naca0006-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 0006 (naca0006-il) Reynolds number: 1,000,000 Max Cl/Cd: 56.63 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca0006-il-1000000.txt Download as CSV file: xf-naca0006-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.7405 0.08222 0.08077 0.0149 1.0000 0.0070
-8.500 -0.7508 0.07216 0.07073 0.0032 1.0000 0.0069
-8.250 -0.7533 0.06306 0.06151 -0.0039 1.0000 0.0068
-8.000 -0.7521 0.05597 0.05427 -0.0074 1.0000 0.0069
-7.750 -0.7462 0.05056 0.04869 -0.0090 1.0000 0.0070
-7.500 -0.7370 0.04574 0.04365 -0.0098 1.0000 0.0072
-7.250 -0.7382 0.03375 0.03109 -0.0092 1.0000 0.0056
-7.000 -0.7345 0.02389 0.02039 -0.0071 1.0000 0.0052
-6.750 -0.7169 0.01969 0.01565 -0.0058 1.0000 0.0053
-6.500 -0.6954 0.01711 0.01268 -0.0048 1.0000 0.0054
-6.250 -0.6709 0.01593 0.01132 -0.0043 1.0000 0.0058
-6.000 -0.6461 0.01480 0.01001 -0.0037 1.0000 0.0061
-5.750 -0.6209 0.01377 0.00881 -0.0031 1.0000 0.0063
-5.500 -0.6001 0.01111 0.00580 -0.0018 1.0000 0.0071
-5.250 -0.5746 0.01047 0.00509 -0.0014 1.0000 0.0079
-5.000 -0.5488 0.00994 0.00451 -0.0010 1.0000 0.0088
-4.750 -0.5225 0.00959 0.00410 -0.0007 1.0000 0.0099
-4.500 -0.4973 0.00878 0.00313 0.0000 1.0000 0.0128
-4.250 -0.4711 0.00839 0.00275 0.0004 1.0000 0.0182
-4.000 -0.4448 0.00812 0.00250 0.0007 1.0000 0.0281
-3.750 -0.4184 0.00796 0.00233 0.0010 1.0000 0.0330
-3.500 -0.3923 0.00774 0.00211 0.0013 1.0000 0.0377
-3.250 -0.3662 0.00761 0.00196 0.0016 1.0000 0.0400
-3.000 -0.3403 0.00737 0.00170 0.0021 1.0000 0.0446
-2.750 -0.3145 0.00718 0.00152 0.0025 1.0000 0.0498
-2.500 -0.2890 0.00689 0.00133 0.0029 1.0000 0.0753
-2.250 -0.2645 0.00622 0.00113 0.0033 1.0000 0.1947
-2.000 -0.2396 0.00570 0.00100 0.0036 1.0000 0.3040
-1.750 -0.2142 0.00530 0.00090 0.0038 1.0000 0.3937
-1.500 -0.1884 0.00495 0.00085 0.0041 1.0000 0.4785
-1.250 -0.1608 0.00464 0.00081 0.0039 0.9995 0.5581
-1.000 -0.1251 0.00426 0.00077 0.0020 0.9958 0.6541
-0.750 -0.0896 0.00395 0.00074 0.0002 0.9904 0.7336
-0.500 -0.0529 0.00365 0.00073 -0.0016 0.9776 0.8123
-0.250 -0.0237 0.00344 0.00073 -0.0015 0.9539 0.8808
0.000 0.0000 0.00338 0.00074 0.0000 0.9238 0.9241
0.250 0.0238 0.00344 0.00073 0.0015 0.8810 0.9537
0.500 0.0528 0.00365 0.00073 0.0016 0.8123 0.9774
0.750 0.0896 0.00396 0.00074 -0.0002 0.7330 0.9905
1.000 0.1251 0.00426 0.00077 -0.0020 0.6539 0.9958
1.250 0.1609 0.00464 0.00081 -0.0039 0.5581 0.9995
1.500 0.1883 0.00495 0.00085 -0.0040 0.4790 1.0000
1.750 0.2140 0.00530 0.00090 -0.0038 0.3938 1.0000
2.000 0.2394 0.00570 0.00100 -0.0035 0.3043 1.0000
2.250 0.2643 0.00622 0.00113 -0.0032 0.1959 1.0000
2.500 0.2888 0.00689 0.00133 -0.0029 0.0756 1.0000
2.750 0.3143 0.00718 0.00152 -0.0024 0.0499 1.0000
3.000 0.3401 0.00737 0.00170 -0.0020 0.0447 1.0000
3.250 0.3660 0.00761 0.00196 -0.0016 0.0400 1.0000
3.500 0.3921 0.00774 0.00211 -0.0012 0.0377 1.0000
3.750 0.4182 0.00796 0.00233 -0.0009 0.0331 1.0000
4.000 0.4446 0.00812 0.00250 -0.0006 0.0281 1.0000
4.250 0.4709 0.00839 0.00274 -0.0003 0.0181 1.0000
4.500 0.4972 0.00878 0.00313 0.0001 0.0127 1.0000
4.750 0.5225 0.00958 0.00409 0.0007 0.0099 1.0000
5.000 0.5487 0.00996 0.00452 0.0010 0.0088 1.0000
5.250 0.5746 0.01046 0.00509 0.0014 0.0079 1.0000
5.500 0.6001 0.01111 0.00581 0.0018 0.0071 1.0000
5.750 0.6211 0.01371 0.00874 0.0031 0.0063 1.0000
6.000 0.6462 0.01479 0.01000 0.0037 0.0061 1.0000
6.250 0.6710 0.01594 0.01133 0.0043 0.0058 1.0000
6.500 0.6954 0.01713 0.01270 0.0048 0.0054 1.0000
6.750 0.7170 0.01971 0.01568 0.0058 0.0053 1.0000
7.000 0.7348 0.02384 0.02034 0.0071 0.0052 1.0000
7.250 0.7383 0.03381 0.03116 0.0091 0.0056 1.0000
10.500 0.7164 0.12067 0.11910 -0.0360 0.0063 1.0000
10.750 0.7174 0.12546 0.12388 -0.0379 0.0063 1.0000
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Polar data table (+)
Polar graphs
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