HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 100,000 Max Cl/Cd: 35.06 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq07-il-100000.txt Download as CSV file: xf-hq07-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/7 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6680 0.08692 0.08246 0.0041 1.0000 0.1042
-8.000 -0.6793 0.08215 0.07774 -0.0006 1.0000 0.1076
-7.750 -0.6925 0.07691 0.07241 -0.0064 1.0000 0.1097
-7.500 -0.7095 0.07305 0.06815 -0.0115 1.0000 0.1111
-6.000 -0.6267 0.03751 0.03037 -0.0120 1.0000 0.0566
-5.750 -0.6064 0.03280 0.02518 -0.0109 1.0000 0.0524
-5.500 -0.5833 0.02891 0.02063 -0.0095 1.0000 0.0507
-5.250 -0.5588 0.02673 0.01795 -0.0084 1.0000 0.0570
-5.000 -0.5339 0.02372 0.01451 -0.0074 1.0000 0.0621
-4.750 -0.5077 0.02172 0.01229 -0.0065 1.0000 0.0696
-4.500 -0.4828 0.01984 0.01038 -0.0057 1.0000 0.0801
-4.250 -0.4584 0.01806 0.00860 -0.0045 1.0000 0.0881
-4.000 -0.4364 0.01653 0.00718 -0.0032 1.0000 0.1045
-3.750 -0.4170 0.01467 0.00589 -0.0019 1.0000 0.1712
-3.500 -0.4133 0.01179 0.00545 0.0029 1.0000 0.6198
-3.250 -0.3986 0.01165 0.00557 0.0073 1.0000 0.7396
-3.000 -0.3842 0.01169 0.00569 0.0120 1.0000 0.8086
-2.750 -0.3684 0.01177 0.00567 0.0166 1.0000 0.8605
-2.500 -0.3375 0.01208 0.00584 0.0187 1.0000 0.9087
-2.250 -0.2791 0.01240 0.00584 0.0143 1.0000 0.9423
-2.000 -0.2189 0.01236 0.00551 0.0079 1.0000 0.9575
-1.750 -0.1578 0.01220 0.00505 0.0010 1.0000 0.9731
-1.500 -0.0839 0.01190 0.00455 -0.0085 1.0000 0.9947
-1.250 -0.0515 0.01155 0.00413 -0.0107 1.0000 1.0000
-1.000 -0.0354 0.01129 0.00385 -0.0096 1.0000 1.0000
-0.750 -0.0206 0.01106 0.00363 -0.0083 1.0000 1.0000
-0.500 -0.0084 0.01087 0.00346 -0.0065 1.0000 1.0000
-0.250 -0.0011 0.01074 0.00336 -0.0038 1.0000 1.0000
0.000 0.0000 0.01068 0.00333 0.0000 1.0000 1.0000
0.250 0.0011 0.01074 0.00336 0.0038 1.0000 1.0000
0.500 0.0084 0.01087 0.00346 0.0065 1.0000 1.0000
0.750 0.0206 0.01106 0.00363 0.0083 1.0000 1.0000
1.000 0.0354 0.01129 0.00385 0.0096 1.0000 1.0000
1.250 0.0515 0.01155 0.00414 0.0107 1.0000 1.0000
1.500 0.0836 0.01190 0.00454 0.0086 0.9947 1.0000
1.750 0.1578 0.01220 0.00505 -0.0010 0.9731 1.0000
2.000 0.2190 0.01236 0.00551 -0.0079 0.9575 1.0000
2.250 0.2791 0.01240 0.00584 -0.0143 0.9423 1.0000
2.500 0.3375 0.01208 0.00584 -0.0187 0.9087 1.0000
2.750 0.3684 0.01177 0.00567 -0.0165 0.8604 1.0000
3.000 0.3842 0.01169 0.00569 -0.0120 0.8086 1.0000
3.250 0.3986 0.01165 0.00557 -0.0073 0.7396 1.0000
3.500 0.4133 0.01179 0.00545 -0.0029 0.6196 1.0000
3.750 0.4169 0.01469 0.00589 0.0019 0.1701 1.0000
4.000 0.4364 0.01653 0.00718 0.0032 0.1046 1.0000
4.250 0.4584 0.01805 0.00860 0.0045 0.0882 1.0000
4.500 0.4828 0.01985 0.01039 0.0057 0.0801 1.0000
4.750 0.5077 0.02173 0.01229 0.0065 0.0696 1.0000
5.000 0.5340 0.02372 0.01451 0.0074 0.0622 1.0000
5.250 0.5589 0.02673 0.01795 0.0084 0.0571 1.0000
5.500 0.5833 0.02892 0.02063 0.0095 0.0506 1.0000
5.750 0.6065 0.03280 0.02518 0.0109 0.0524 1.0000
6.000 0.6268 0.03747 0.03034 0.0120 0.0566 1.0000
6.250 0.6545 0.05318 0.04838 0.0089 0.1842 1.0000
6.500 0.6675 0.05662 0.05184 0.0096 0.1672 1.0000
6.750 0.6797 0.06033 0.05555 0.0103 0.1526 1.0000
8.000 0.7102 0.08108 0.07643 0.0090 0.0977 1.0000
8.250 0.6931 0.08551 0.08103 0.0038 0.0971 1.0000
8.500 0.6806 0.09047 0.08599 -0.0011 0.0968 1.0000
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Polar data table (+)
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