HQ 2.0/8 AIRFOIL (hq208-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.0/8 AIRFOIL (hq208-il) Reynolds number: 1,000,000 Max Cl/Cd: 112.75 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq208-il-1000000.txt Download as CSV file: xf-hq208-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.0/8 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.3890 0.09913 0.09758 -0.0194 1.0000 0.0066
-10.000 -0.3886 0.09475 0.09321 -0.0208 1.0000 0.0067
-9.750 -0.3879 0.09053 0.08900 -0.0221 1.0000 0.0069
-9.500 -0.3868 0.08641 0.08489 -0.0233 1.0000 0.0069
-9.250 -0.3864 0.08219 0.08068 -0.0246 1.0000 0.0070
-9.000 -0.3861 0.07809 0.07660 -0.0259 1.0000 0.0071
-8.750 -0.3867 0.07394 0.07245 -0.0272 1.0000 0.0072
-8.500 -0.3886 0.06962 0.06816 -0.0286 1.0000 0.0072
-8.250 -0.3924 0.06508 0.06364 -0.0302 1.0000 0.0073
-8.000 -0.3973 0.06068 0.05925 -0.0319 1.0000 0.0073
-7.750 -0.4063 0.05611 0.05471 -0.0338 1.0000 0.0073
-7.500 -0.4259 0.05073 0.04938 -0.0378 1.0000 0.0072
-7.250 -0.4393 0.04484 0.04345 -0.0419 1.0000 0.0071
-7.000 -0.4275 0.03766 0.03613 -0.0490 0.9959 0.0072
-6.750 -0.4098 0.03149 0.02980 -0.0544 0.9908 0.0074
-6.500 -0.3884 0.02597 0.02408 -0.0587 0.9861 0.0076
-6.250 -0.3640 0.02094 0.01882 -0.0622 0.9820 0.0080
-6.000 -0.3421 0.01675 0.01437 -0.0640 0.9733 0.0084
-5.250 -0.2841 0.01711 0.01324 -0.0669 0.9655 0.0080
-5.000 -0.2565 0.01347 0.00908 -0.0660 0.9567 0.0060
-4.500 -0.2066 0.01005 0.00515 -0.0648 0.9375 0.0076
-4.250 -0.1806 0.00952 0.00454 -0.0644 0.9285 0.0083
-4.000 -0.1546 0.00898 0.00391 -0.0639 0.9193 0.0088
-3.750 -0.1282 0.00862 0.00348 -0.0636 0.9111 0.0097
-3.500 -0.1018 0.00820 0.00295 -0.0632 0.9027 0.0104
-3.250 -0.0750 0.00778 0.00241 -0.0628 0.8945 0.0109
-3.000 -0.0482 0.00735 0.00181 -0.0624 0.8868 0.0144
-2.750 -0.0212 0.00700 0.00154 -0.0622 0.8784 0.0406
-2.250 0.0325 0.00644 0.00124 -0.0619 0.8620 0.1382
-2.000 0.0586 0.00590 0.00109 -0.0619 0.8525 0.2678
-1.750 0.0830 0.00505 0.00096 -0.0617 0.8424 0.5101
-1.500 0.1096 0.00488 0.00091 -0.0614 0.8323 0.5779
-1.250 0.1370 0.00480 0.00088 -0.0613 0.8223 0.6145
-1.000 0.1645 0.00475 0.00085 -0.0611 0.8126 0.6412
-0.750 0.1918 0.00471 0.00082 -0.0609 0.8023 0.6657
-0.500 0.2190 0.00468 0.00081 -0.0607 0.7916 0.6940
-0.250 0.2458 0.00463 0.00083 -0.0604 0.7796 0.7285
0.000 0.2731 0.00462 0.00083 -0.0601 0.7679 0.7509
0.250 0.3003 0.00461 0.00084 -0.0599 0.7570 0.7697
0.500 0.3275 0.00461 0.00085 -0.0597 0.7452 0.7878
0.750 0.3547 0.00462 0.00086 -0.0595 0.7320 0.8015
1.000 0.3819 0.00464 0.00087 -0.0593 0.7186 0.8150
1.250 0.4088 0.00466 0.00089 -0.0590 0.7033 0.8302
1.500 0.4353 0.00468 0.00093 -0.0587 0.6866 0.8477
1.750 0.4612 0.00469 0.00096 -0.0582 0.6688 0.8710
2.000 0.4854 0.00465 0.00100 -0.0572 0.6484 0.9122
2.250 0.5229 0.00467 0.00103 -0.0593 0.6213 1.0000
2.500 0.5491 0.00487 0.00110 -0.0590 0.5868 1.0000
2.750 0.5737 0.00525 0.00122 -0.0585 0.5182 1.0000
3.000 0.5975 0.00576 0.00138 -0.0579 0.4370 1.0000
3.250 0.6227 0.00613 0.00154 -0.0576 0.3841 1.0000
3.500 0.6478 0.00653 0.00173 -0.0572 0.3300 1.0000
3.750 0.6724 0.00700 0.00194 -0.0568 0.2736 1.0000
4.000 0.6981 0.00734 0.00216 -0.0566 0.2402 1.0000
4.250 0.7240 0.00764 0.00235 -0.0563 0.2122 1.0000
4.500 0.7488 0.00807 0.00258 -0.0560 0.1667 1.0000
4.750 0.7729 0.00861 0.00287 -0.0555 0.1162 1.0000
5.000 0.7967 0.00920 0.00323 -0.0550 0.0711 1.0000
5.250 0.8213 0.00967 0.00358 -0.0546 0.0459 1.0000
5.500 0.8468 0.01000 0.00387 -0.0544 0.0346 1.0000
5.750 0.8722 0.01035 0.00416 -0.0540 0.0242 1.0000
6.000 0.8965 0.01088 0.00460 -0.0535 0.0098 1.0000
6.250 0.9212 0.01133 0.00504 -0.0530 0.0057 1.0000
6.500 0.9462 0.01174 0.00551 -0.0526 0.0050 1.0000
6.750 0.9709 0.01219 0.00601 -0.0521 0.0046 1.0000
7.000 0.9950 0.01272 0.00665 -0.0515 0.0042 1.0000
7.250 1.0180 0.01342 0.00745 -0.0507 0.0037 1.0000
7.500 1.0368 0.01480 0.00905 -0.0491 0.0032 1.0000
7.750 1.0593 0.01553 0.00988 -0.0483 0.0030 1.0000
8.000 1.0796 0.01657 0.01106 -0.0471 0.0030 1.0000
8.250 1.0983 0.01789 0.01255 -0.0457 0.0029 1.0000
8.500 1.1160 0.01942 0.01426 -0.0442 0.0029 1.0000
8.750 1.1316 0.02138 0.01645 -0.0423 0.0029 1.0000
9.000 1.1452 0.02377 0.01911 -0.0404 0.0030 1.0000
9.250 1.1510 0.02775 0.02346 -0.0375 0.0033 1.0000
9.500 1.1688 0.02876 0.02461 -0.0363 0.0035 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.0/8 AIRFOIL (hq208-il)