HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 200,000 Max Cl/Cd: 43.14 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq07-il-200000.txt Download as CSV file: xf-hq07-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5805 0.09103 0.08791 0.0073 1.0000 0.0367
-9.000 -0.5822 0.08619 0.08310 0.0056 1.0000 0.0374
-8.000 -0.7065 0.06777 0.06425 -0.0123 1.0000 0.0350
-7.750 -0.7013 0.06391 0.06038 -0.0124 1.0000 0.0359
-7.500 -0.6960 0.06015 0.05655 -0.0128 1.0000 0.0369
-7.250 -0.6897 0.05619 0.05245 -0.0133 1.0000 0.0381
-7.000 -0.6816 0.05203 0.04810 -0.0136 1.0000 0.0396
-6.750 -0.6711 0.04780 0.04358 -0.0137 1.0000 0.0416
-6.500 -0.6623 0.04394 0.03890 -0.0130 1.0000 0.0472
-6.250 -0.6402 0.03442 0.02886 -0.0106 1.0000 0.0211
-6.000 -0.6241 0.02965 0.02361 -0.0095 1.0000 0.0199
-5.750 -0.6035 0.02585 0.01928 -0.0082 1.0000 0.0196
-5.500 -0.5800 0.02295 0.01589 -0.0070 1.0000 0.0206
-5.250 -0.5575 0.02051 0.01315 -0.0062 1.0000 0.0256
-5.000 -0.5324 0.01889 0.01124 -0.0054 1.0000 0.0302
-4.750 -0.5082 0.01680 0.00898 -0.0043 1.0000 0.0373
-4.500 -0.4832 0.01587 0.00794 -0.0036 1.0000 0.0451
-4.250 -0.4602 0.01451 0.00661 -0.0028 1.0000 0.0523
-4.000 -0.4360 0.01363 0.00563 -0.0019 1.0000 0.0586
-3.750 -0.4134 0.01246 0.00450 -0.0010 1.0000 0.0737
-3.500 -0.3946 0.01059 0.00349 0.0000 1.0000 0.2274
-3.250 -0.3795 0.00891 0.00310 0.0017 1.0000 0.5024
-3.000 -0.3611 0.00837 0.00311 0.0039 1.0000 0.6498
-2.750 -0.3407 0.00817 0.00303 0.0059 1.0000 0.7188
-2.500 -0.3208 0.00805 0.00302 0.0081 1.0000 0.7691
-2.250 -0.3012 0.00798 0.00300 0.0103 1.0000 0.8069
-2.000 -0.2834 0.00790 0.00299 0.0130 1.0000 0.8405
-1.750 -0.2672 0.00786 0.00299 0.0159 1.0000 0.8738
-1.500 -0.2493 0.00784 0.00300 0.0183 1.0000 0.9012
-1.250 -0.2246 0.00786 0.00296 0.0191 1.0000 0.9228
-1.000 -0.1905 0.00793 0.00299 0.0177 1.0000 0.9449
-0.750 -0.1455 0.00802 0.00303 0.0141 1.0000 0.9637
-0.500 -0.0946 0.00810 0.00306 0.0092 1.0000 0.9784
-0.250 -0.0420 0.00812 0.00304 0.0038 1.0000 0.9893
0.000 -0.0001 0.00812 0.00303 0.0000 1.0000 1.0000
0.250 0.0420 0.00812 0.00304 -0.0038 0.9893 1.0000
0.500 0.0948 0.00810 0.00306 -0.0093 0.9784 1.0000
0.750 0.1456 0.00802 0.00303 -0.0142 0.9636 1.0000
1.000 0.1906 0.00793 0.00299 -0.0178 0.9449 1.0000
1.250 0.2246 0.00786 0.00296 -0.0191 0.9228 1.0000
1.500 0.2494 0.00784 0.00300 -0.0183 0.9012 1.0000
1.750 0.2671 0.00786 0.00299 -0.0159 0.8737 1.0000
2.000 0.2833 0.00790 0.00299 -0.0129 0.8405 1.0000
2.250 0.3012 0.00798 0.00300 -0.0103 0.8070 1.0000
2.500 0.3208 0.00805 0.00302 -0.0081 0.7690 1.0000
2.750 0.3406 0.00817 0.00303 -0.0059 0.7186 1.0000
3.000 0.3611 0.00837 0.00310 -0.0039 0.6498 1.0000
3.250 0.3795 0.00891 0.00310 -0.0017 0.5023 1.0000
3.500 0.3945 0.01059 0.00349 0.0000 0.2265 1.0000
3.750 0.4134 0.01247 0.00451 0.0010 0.0735 1.0000
4.000 0.4359 0.01363 0.00563 0.0019 0.0587 1.0000
4.250 0.4602 0.01451 0.00661 0.0028 0.0524 1.0000
4.500 0.4832 0.01587 0.00795 0.0036 0.0451 1.0000
4.750 0.5083 0.01679 0.00897 0.0043 0.0373 1.0000
5.000 0.5324 0.01889 0.01124 0.0054 0.0303 1.0000
5.250 0.5576 0.02064 0.01330 0.0063 0.0261 1.0000
5.500 0.5801 0.02294 0.01588 0.0070 0.0205 1.0000
5.750 0.6035 0.02586 0.01928 0.0082 0.0196 1.0000
6.000 0.6242 0.02963 0.02358 0.0095 0.0199 1.0000
6.250 0.6402 0.03444 0.02889 0.0106 0.0211 1.0000
7.250 0.6478 0.04697 0.04359 0.0132 0.0397 1.0000
7.500 0.6465 0.05211 0.04888 0.0123 0.0386 1.0000
7.750 0.6406 0.05729 0.05418 0.0111 0.0379 1.0000
8.000 0.6273 0.06239 0.05936 0.0093 0.0377 1.0000
8.250 0.6053 0.06892 0.06591 0.0038 0.0384 1.0000
8.500 0.5933 0.07525 0.07221 -0.0005 0.0383 1.0000
8.750 0.5865 0.08092 0.07785 -0.0035 0.0380 1.0000
9.000 0.5823 0.08611 0.08301 -0.0057 0.0373 1.0000
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