HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 1,000,000 Max Cl/Cd: 59.14 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq07-il-1000000.txt Download as CSV file: xf-hq07-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.5855 0.14723 0.14576 0.0361 1.0000 0.0060
-12.250 -0.5827 0.14330 0.14183 0.0341 1.0000 0.0060
-6.250 -0.6618 0.01880 0.01457 -0.0059 1.0000 0.0046
-6.000 -0.6429 0.01439 0.00960 -0.0039 1.0000 0.0040
-5.750 -0.6211 0.01221 0.00711 -0.0026 1.0000 0.0038
-5.500 -0.5976 0.01097 0.00568 -0.0017 1.0000 0.0039
-5.250 -0.5731 0.01011 0.00468 -0.0010 1.0000 0.0041
-5.000 -0.5480 0.00947 0.00393 -0.0004 1.0000 0.0046
-4.750 -0.5228 0.00883 0.00328 0.0002 1.0000 0.0108
-4.500 -0.4965 0.00867 0.00307 0.0004 1.0000 0.0141
-4.250 -0.4707 0.00837 0.00276 0.0008 1.0000 0.0174
-4.000 -0.4449 0.00816 0.00255 0.0011 1.0000 0.0199
-3.750 -0.4190 0.00804 0.00244 0.0014 1.0000 0.0219
-3.500 -0.3944 0.00764 0.00205 0.0020 1.0000 0.0319
-3.250 -0.3703 0.00724 0.00178 0.0027 1.0000 0.0634
-3.000 -0.3377 0.00674 0.00151 0.0013 0.9947 0.1265
-2.750 -0.3028 0.00621 0.00132 -0.0007 0.9853 0.2190
-2.500 -0.2672 0.00571 0.00114 -0.0028 0.9725 0.3059
-2.250 -0.2348 0.00516 0.00098 -0.0041 0.9539 0.4291
-2.000 -0.2098 0.00466 0.00087 -0.0037 0.9292 0.5530
-1.750 -0.1853 0.00443 0.00081 -0.0028 0.9057 0.6300
-1.500 -0.1599 0.00436 0.00077 -0.0022 0.8852 0.6672
-1.250 -0.1338 0.00432 0.00075 -0.0017 0.8675 0.6981
-1.000 -0.1074 0.00428 0.00073 -0.0013 0.8526 0.7227
-0.750 -0.0810 0.00426 0.00072 -0.0008 0.8361 0.7448
-0.500 -0.0542 0.00427 0.00069 -0.0005 0.8176 0.7589
-0.250 -0.0271 0.00427 0.00068 -0.0002 0.8010 0.7722
0.000 0.0000 0.00427 0.00068 0.0000 0.7868 0.7867
0.250 0.0272 0.00427 0.00068 0.0002 0.7722 0.8010
0.500 0.0542 0.00427 0.00069 0.0005 0.7588 0.8178
0.750 0.0811 0.00426 0.00072 0.0008 0.7448 0.8360
1.000 0.1075 0.00428 0.00073 0.0012 0.7227 0.8526
1.250 0.1339 0.00432 0.00075 0.0017 0.6980 0.8675
1.500 0.1600 0.00436 0.00077 0.0022 0.6672 0.8852
1.750 0.1854 0.00443 0.00081 0.0028 0.6300 0.9058
2.000 0.2098 0.00466 0.00087 0.0037 0.5508 0.9293
2.250 0.2348 0.00516 0.00098 0.0041 0.4274 0.9539
2.500 0.2672 0.00571 0.00114 0.0028 0.3060 0.9727
2.750 0.3027 0.00621 0.00132 0.0007 0.2183 0.9854
3.000 0.3378 0.00674 0.00151 -0.0013 0.1263 0.9947
3.250 0.3702 0.00724 0.00178 -0.0027 0.0634 1.0000
3.500 0.3943 0.00764 0.00205 -0.0020 0.0318 1.0000
3.750 0.4189 0.00804 0.00245 -0.0014 0.0219 1.0000
4.000 0.4448 0.00816 0.00255 -0.0011 0.0200 1.0000
4.250 0.4707 0.00837 0.00277 -0.0007 0.0175 1.0000
4.500 0.4964 0.00867 0.00308 -0.0004 0.0141 1.0000
4.750 0.5228 0.00884 0.00329 -0.0002 0.0109 1.0000
5.000 0.5480 0.00947 0.00393 0.0004 0.0046 1.0000
5.250 0.5731 0.01011 0.00467 0.0010 0.0041 1.0000
5.500 0.5976 0.01098 0.00569 0.0017 0.0039 1.0000
5.750 0.6211 0.01220 0.00710 0.0026 0.0038 1.0000
6.000 0.6430 0.01433 0.00953 0.0039 0.0040 1.0000
6.250 0.6618 0.01884 0.01462 0.0059 0.0047 1.0000
14.000 0.7175 0.17531 0.17362 -0.0485 0.0059 1.0000
14.250 0.7241 0.17936 0.17766 -0.0505 0.0059 1.0000
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