HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il) Reynolds number: 100,000 Max Cl/Cd: 51.39 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1585-il-100000.txt Download as CSV file: xf-hq1585-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8.5 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5424 0.09030 0.08568 -0.0240 1.0000 0.1185
-8.250 -0.5133 0.08647 0.08176 -0.0183 1.0000 0.1247
-8.000 -0.5233 0.08291 0.07830 -0.0218 1.0000 0.1303
-7.500 -0.5226 0.07501 0.07051 -0.0253 1.0000 0.1413
-7.250 -0.5335 0.06964 0.06513 -0.0317 1.0000 0.1489
-7.000 -0.5407 0.06575 0.06110 -0.0363 1.0000 0.1612
-6.750 -0.5181 0.06312 0.05863 -0.0315 1.0000 0.1685
-6.500 -0.5131 0.05983 0.05533 -0.0319 1.0000 0.1833
-6.000 -0.4993 0.04010 0.03379 -0.0420 1.0000 0.0898
-5.750 -0.4830 0.03532 0.02822 -0.0406 1.0000 0.0769
-5.500 -0.4667 0.03160 0.02426 -0.0394 1.0000 0.0737
-5.250 -0.4485 0.02845 0.02061 -0.0378 1.0000 0.0705
-5.000 -0.4290 0.02600 0.01767 -0.0362 1.0000 0.0694
-4.750 -0.4096 0.02443 0.01585 -0.0347 1.0000 0.0733
-4.500 -0.3889 0.02290 0.01399 -0.0332 1.0000 0.0759
-4.250 -0.3673 0.02152 0.01231 -0.0317 1.0000 0.0777
-4.000 -0.3462 0.01975 0.01055 -0.0305 1.0000 0.0812
-3.750 -0.3256 0.01883 0.00960 -0.0293 1.0000 0.0898
-3.500 -0.3051 0.01765 0.00854 -0.0281 1.0000 0.0984
-3.250 -0.2843 0.01666 0.00762 -0.0270 1.0000 0.1124
-3.000 -0.2628 0.01553 0.00680 -0.0262 1.0000 0.1603
-2.750 -0.2508 0.01305 0.00670 -0.0234 1.0000 0.6092
-2.500 -0.2391 0.01313 0.00702 -0.0193 1.0000 0.7186
-2.250 -0.2275 0.01321 0.00714 -0.0154 1.0000 0.7828
-2.000 -0.2176 0.01317 0.00715 -0.0110 1.0000 0.8359
-1.750 -0.2079 0.01306 0.00713 -0.0063 1.0000 0.8988
-1.500 -0.1401 0.01311 0.00701 -0.0131 1.0000 0.9899
-1.250 -0.1224 0.01303 0.00675 -0.0134 1.0000 1.0000
-1.000 -0.1029 0.01308 0.00662 -0.0136 1.0000 1.0000
-0.750 -0.0808 0.01324 0.00660 -0.0140 1.0000 1.0000
-0.500 -0.0405 0.01363 0.00679 -0.0177 0.9944 1.0000
-0.250 0.0052 0.01402 0.00704 -0.0222 0.9855 1.0000
0.000 0.0492 0.01440 0.00730 -0.0262 0.9765 1.0000
0.250 0.0973 0.01482 0.00762 -0.0309 0.9686 1.0000
0.500 0.1376 0.01510 0.00784 -0.0340 0.9582 1.0000
0.750 0.1833 0.01535 0.00805 -0.0379 0.9466 1.0000
1.000 0.2314 0.01551 0.00820 -0.0419 0.9346 1.0000
1.250 0.2818 0.01560 0.00831 -0.0463 0.9238 1.0000
1.500 0.3393 0.01547 0.00826 -0.0516 0.9127 1.0000
1.750 0.3830 0.01536 0.00821 -0.0542 0.8993 1.0000
2.000 0.4208 0.01531 0.00822 -0.0556 0.8860 1.0000
2.250 0.4555 0.01526 0.00827 -0.0563 0.8723 1.0000
2.500 0.4883 0.01516 0.00825 -0.0564 0.8579 1.0000
2.750 0.5192 0.01503 0.00821 -0.0560 0.8428 1.0000
3.000 0.5465 0.01493 0.00819 -0.0549 0.8253 1.0000
3.250 0.5720 0.01480 0.00817 -0.0533 0.8051 1.0000
3.500 0.5977 0.01454 0.00797 -0.0513 0.7827 1.0000
3.750 0.6220 0.01429 0.00776 -0.0491 0.7569 1.0000
4.000 0.6445 0.01407 0.00762 -0.0468 0.7240 1.0000
4.250 0.6667 0.01373 0.00732 -0.0442 0.6796 1.0000
4.500 0.6883 0.01355 0.00697 -0.0416 0.6174 1.0000
4.750 0.7086 0.01379 0.00692 -0.0392 0.5366 1.0000
5.000 0.7285 0.01440 0.00713 -0.0373 0.4560 1.0000
5.250 0.7484 0.01521 0.00764 -0.0358 0.3877 1.0000
5.500 0.7676 0.01614 0.00826 -0.0344 0.3185 1.0000
5.750 0.7830 0.01760 0.00915 -0.0327 0.2123 1.0000
6.000 0.7976 0.01962 0.01045 -0.0308 0.1380 1.0000
6.250 0.8182 0.02110 0.01183 -0.0295 0.1177 1.0000
6.500 0.8406 0.02268 0.01335 -0.0285 0.1069 1.0000
6.750 0.8639 0.02432 0.01494 -0.0277 0.0961 1.0000
7.000 0.8875 0.02619 0.01673 -0.0272 0.0848 1.0000
7.250 0.9095 0.02766 0.01831 -0.0264 0.0727 1.0000
7.500 0.9327 0.02973 0.02070 -0.0253 0.0636 1.0000
7.750 0.9541 0.03250 0.02369 -0.0243 0.0563 1.0000
8.000 0.9751 0.03550 0.02711 -0.0230 0.0529 1.0000
8.250 0.9938 0.03879 0.03077 -0.0217 0.0510 1.0000
8.500 1.0093 0.04234 0.03464 -0.0206 0.0491 1.0000
8.750 1.0169 0.04811 0.04079 -0.0194 0.0477 1.0000
9.000 1.0211 0.05111 0.04438 -0.0172 0.0468 1.0000
9.250 1.0228 0.05567 0.04935 -0.0155 0.0470 1.0000
9.500 1.0274 0.06069 0.05465 -0.0143 0.0480 1.0000
9.750 0.9900 0.06590 0.06081 -0.0109 0.0525 1.0000
10.000 0.9644 0.07083 0.06598 -0.0094 0.0542 1.0000
10.250 0.9403 0.07577 0.07106 -0.0098 0.0553 1.0000
10.500 0.9177 0.08149 0.07689 -0.0121 0.0564 1.0000
10.750 0.9057 0.08738 0.08281 -0.0145 0.0580 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/8.5 AIRFOIL (hq1585-il)